Hybrid Propulsion System for CubeSat Applicationsepubs.surrey.ac.uk/812899/1/PhD_Thesis_AOD.pdf ·...

Hybrid Propulsion System for CubeSat

Applications

Ahmed Ozomata David

Submitted for the Degree of

Doctor of Philosophy

from the University of Surrey

Surrey Space Centre

Department of Electronic Engineering

Faculty of Engineering and Physical Sciences

University of Surrey

Guildford, Surrey, GU2 7XH, UK.

September 2016

cAhmed Ozomata David 2016

Abstract

The CubeSats platform has become a common basis for the development and flight of very

small, low cost spacecraft-particularly amongst University groups. The smallest CubeSats

are just 1 litre in volume-comprising a 10 𝑐𝑚 x 10 𝑐𝑚 x 10 𝑐𝑚 unit-“1𝑈”. Multiples of this

unit are also flown: 2𝑈 and 3𝑈 (which fit the standard launch “pod”) and, at the larger

scale, 6𝑈 , 12𝑈 and potentially 27𝑈 . The spacecraft generally do not carry propulsion

systems and so their orbit is dictated by the initial orbital injection from the launch

vehicle. This research aims at producing a novel chemical micropropulsion system based

on a mixture of sodium hydroxide and water (the oxidiser) and aluminium (the fuel)

suitable for CubeSats. The choice of the propellants was based on the availability and

cost of materials; long storage without degrading; moderate temperature and exothermic

reaction without any thermal control threat to the microsatellite structure; high energy

density per unit volume for the volume constraint satellite; and the propulsion system

will require minimal power from the CubeSat electrical bus system. Initial experimental

findings revealed that oxidiser of 12.50𝑚𝑜𝑙/𝑘𝑔 molar concentration produced the fastest

reaction rate, and a reaction of 6 𝑔 of fuel to 3 𝑔 of oxidiser produced a peak performance

of 0.032𝑁 thrust and 45 𝑠 specific impulse. Multiple injections of the oxidiser for repeat

cycles were also demonstrated with different fuel to oxidiser ratios. The energy utilisation

of the propulsion system was calculated and it revealed that about 98% of the exhaust

was water vapour , while only about 2% was hydrogen gas. It was also found out that

about 4% of the total generated enthalpy was converted into useful thrust, while the

remaining percentage was used by about 98% of the injected water to change phase from

liquid to gas. This result assumed that the reaction volume was essentially adiabatic.

Though the specific impulse of the propellant is moderate, the thruster is capable of

delivering a ΔV of about 57𝑚/𝑠 to a 1U CubeSat of 1.33 𝑘𝑔. However, one of the

drawbacks of the system is that the firing time is about 1 𝑚𝑖𝑛𝑢𝑡𝑒 after the injection of

oxidiser, making this system inappropriate for attitude control purposes.

ii

Acknowledgements

This enduring journey has been made possible by this set of special people that their

contributions toward the completion of this program can not just be mentioned by words.

Howbeit, let me use this privilege to acknowledge my Dad and Mum, Mr Amodu Oyibo

Lawal and Mrs Ayisetu Mariya Amodu, for your unending love, care and prayers. I am

deeply grateful. And to all my siblings especially Mr Oyibo Sunday Amodu and late

Mr Adeku Joseph Amodu, I say thank you. You saw this potential in me and did not

let go of it at that tiny age, and now this is it! A million thanks to my lovely wife and

children: Oziohu Glory and Adinoyi, Onimisi, Onize and Adavize. The smiles on your

faces during this sacrifice kept me going even when the journey seemed tough. We share

the research story together. Thanks also to my Uncle, Mr M.A. Momoh and Mummy,

late Mrs Ester Momoh. You took me as I was and gave me the opportunity to discover

myself. Am grateful. Thanks also to Mr Nathaniel Salawu and family for your valuable

advice and encouragement. To the staff and management of Nigerian Communication

Satellite Limited, you made it possible for me to achieve my dream, and so thank you

very much. To the financier of the PhD program, Petroleum Technology Development

Fund (PTDF) of Nigeria, I say thank you. Million thanks to my supervisors, Dr Aaron

Knoll and Prof. Phil Palmer for your unflinching supports and in-depth contributions

during the course of this program. Thanks also to all the SSC propulsion group-both

formal and the present members for the propulsion engineering and update discussions.

Thanks to the three Toms, Charlie, Andrea, Max, Antonio, Gebi and Ahmad for your

contributions. A very big thank you to all the administrative and technical staff of SSC:

Karen, Louise, Andy and David, for your supports and making my stay at Surrey a

very conducive one. Thanks also to my colleague: Yusuf, Pam, Ugah, Mahmoud and

Modibbo for our quality time together. Thanks to Dr Ibrahim, Dr Ikpaya, Dr Daji, Dr

Tanko and Dr Okonor for your valuable advice and supports during this program. Above

all, I give all thanks to God Almighty for His love, grace and mercy.

iii

Contents

Abstract ii

Acknowledgements iii

List of Figures vii

List of Tables xi

Nomenclature xii

1 Introduction 1

1.1 CubeSat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.1.1 CubeSats Flown with Propulsion Systems . . . . . . . . . . . . . 3

1.2 Overview of Micropropulsion Systems . . . . . . . . . . . . . . . . . . . 6

1.2.1 Electric Micropropulsion Systems . . . . . . . . . . . . . . . . . . 6

1.2.1.1 Resistojets . . . . . . . . . . . . . . . . . . . . . . . . . . 7

1.2.1.2 Arcjets . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

1.2.1.3 Microcavity Discharge Thruster . . . . . . . . . . . . . 9

1.2.1.4 Ion Engines . . . . . . . . . . . . . . . . . . . . . . . . . 10

1.2.1.5 Hall Thrusters . . . . . . . . . . . . . . . . . . . . . . . . 11

1.2.1.6 Micro Pulse Plasma Thrusters . . . . . . . . . . . . . . 13

1.2.1.7 Micro Laser Ablation Thruster . . . . . . . . . . . . . . 14

1.2.1.8 Vacuum Arc Thruster . . . . . . . . . . . . . . . . . . . 15

1.2.1.9 Field Emission Electric Propulsion . . . . . . . . . . . . 16

1.2.2 Chemical Micropropulsion Systems . . . . . . . . . . . . . . . . . . 17

1.2.2.1 Cold Gas Thruster . . . . . . . . . . . . . . . . . . . . . . 17

1.2.2.2 Warm Gas Thruster . . . . . . . . . . . . . . . . . . . . 20

iv

1.2.2.3 Monopropellant Systems . . . . . . . . . . . . . . . . . . 21

1.2.2.4 Bipropellant Thrusters . . . . . . . . . . . . . . . . . . 22

1.2.2.5 Solid Rocket Motor . . . . . . . . . . . . . . . . . . . . 23

1.2.2.6 Hybrid Propulsion System . . . . . . . . . . . . . . . . 25

1.3 CubeSat Requirements for Propulsion System . . . . . . . . . . . . . . . 25

1.3.1 Propulsion Requirements for CubeSat Missions . . . . . . . . . . 26

1.3.2 Chemical Propulsion Trade-off . . . . . . . . . . . . . . . . . . . 28

1.3.2.1 Overview of Hybrid Rocket Motor . . . . . . . . . . . . 28

1.4 Motivation and Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . 32

1.5 Novelty and Research Achievements . . . . . . . . . . . . . . . . . . . . 33

2 Theory 34

2.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

2.2 Aluminium, Sodium Hydroxide, Water Oxidation Reaction . . . . . . . 36

2.3 Thermodynamics and Gas Dynamics . . . . . . . . . . . . . . . . . . . . 40

2.4 Theoretical Performance Analysis of the Propulsion System . . . . . . . 45

2.4.1 Thruster Design . . . . . . . . . . . . . . . . . . . . . . . . . . . 49

2.4.1.1 Reaction Chamber . . . . . . . . . . . . . . . . . . . . . 49

2.4.1.2 Nozzle . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

3 Experimental Setup 52

3.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

3.2 Oxidizer Feed System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

3.3 Data Acquisition System . . . . . . . . . . . . . . . . . . . . . . . . . . . 54

3.3.1 Temperature and Pressure Sensors . . . . . . . . . . . . . . . . . 54

3.3.2 DAQ Measurement Hardware . . . . . . . . . . . . . . . . . . . . 55

3.3.3 LabVIEW Software . . . . . . . . . . . . . . . . . . . . . . . . . 56

3.4 Vacuum Facilities and Thrust Balance . . . . . . . . . . . . . . . . . . . . 57

3.4.1 The Pagasus Vacuum Chamber . . . . . . . . . . . . . . . . . . . . 57

3.4.2 Thrust Balance Arrangement . . . . . . . . . . . . . . . . . . . . 58

3.4.3 Calibration and Data Analysis . . . . . . . . . . . . . . . . . . . 59

3.5 Complete Experimental Setup . . . . . . . . . . . . . . . . . . . . . . . . 62

4 Results and Discussion 64

4.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64

v

4.2 Stages of Experiments . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64

4.2.1 Reaction Chemistry of the Propellants at Ambient Conditions . 64

4.2.2 Temperature and Pressure Rise in a control volume under Vacuum

Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

4.2.3 Effect of Oxidiser Concentration on Thruster Characterisation . 68

4.2.4 Fuel/Oxidiser Ratio on Thruster Performance . . . . . . . . . . . 72

4.2.5 Propellant Mass Effect on Thruster Performance . . . . . . . . . 74

4.2.6 Impact of Repeat Cycles on Thruster Performance . . . . . . . . 76

4.2.7 Effect of Nozzle throat Diameter on Thruster Performance . . . 78

4.3 Reaction Pattern of the Propulsion System . . . . . . . . . . . . . . . . 80

4.4 Energy Conversion Efficiency of the Propulsion System . . . . . . . . . . . 81

4.4.1 Chemical Analysis of the Residual Propellants . . . . . . . . . . . 87

4.5 Comparison Between Design Target, Theoretical and Prototype Performances 89

4.6 Comparison with the State-of-the-Art . . . . . . . . . . . . . . . . . . . . 91

4.7 Summary of Experimental Findings . . . . . . . . . . . . . . . . . . . . 92

4.7.0.1 Proposed Mechanical Design of the Hybrid Propulsion

System . . . . . . . . . . . . . . . . . . . . . . . . . . . 93

5 Conclusions and Future Work 96

5.1 Novelty and Research Achievements . . . . . . . . . . . . . . . . . . . . 98

5.2 Future Work . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99

References 101

Appendix A Detail Drawings of the Hybrid Propulsion Thruster 117

Appendix B Experiment Hardware 122

Appendix C Program Codes 136

C.1 Solenoid Valves Control Program . . . . . . . . . . . . . . . . . . . . . . 136

C.2 Thrust Balance Calibration Constant Program . . . . . . . . . . . . . . 138

C.3 Thrust Response Program of One-shot Experiment . . . . . . . . . . . . 142

C.4 Thrust Response Program of Repeat Cycle Injection . . . . . . . . . . . 143

vi

List of Figures

1.1 1U cube satellite [5] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.2 Concept of constellation flight [17] . . . . . . . . . . . . . . . . . . . . . 2

1.3 CanX-2 NANOPS system [20] . . . . . . . . . . . . . . . . . . . . . . . 3

1.4 STRaND-1 propulsion systems: (a) Butane resistojet (b) Pulsed plasma

thruster [25] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.5 Delfi-n3Xt cold gas generator thruster components [29] . . . . . . . . . . 5

1.6 Schematic and complete views of vaporizing liquid thruster . . . . . . . . 7

1.7 3-Watt CMOS resistojet on a 2-micron TinyChip die layout [40] . . . . 8

1.8 Schematic of an arcjet thruster [44] . . . . . . . . . . . . . . . . . . . . 9

1.9 Schematic of an insulated electrodes of a microcavity thruster [47] . . . 10

1.10 Schematic diagrams of a 3D view and a longitudinal cross-section view of

the ion thruster developed at Pennsylvania State University [48] . . . . . 11

1.11 Schematic diagram of an SPT Hall thruster, showing the electrodes and

the radial magnetic field [49] . . . . . . . . . . . . . . . . . . . . . . . . 12

1.12 Schematic diagram and photo shot of a low-power miniaturised Hall

thruster (TCHT-4) [50] . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

1.13 𝜇PPT for concepts for microsatellites . . . . . . . . . . . . . . . . . . . . 14

1.14 Micro laser ablation thruster concept [57] . . . . . . . . . . . . . . . . . 14

1.15 Schematic diagram of a magnetically enhanced vacuum arc thruster [62] 15

1.16 Schematic diagram of a field emission electric propulsion [65] . . . . . . 16

1.17 Schematic view of cold gas thruster [38] . . . . . . . . . . . . . . . . . . 18

1.18 Schematic diagram of MPS-110 cold gas thruster developing by Aerojet [69] 19

1.19 Gas generator cartridges [72] . . . . . . . . . . . . . . . . . . . . . . . . 20

1.20 Schematic of a novel warm gas propulsion system [74] . . . . . . . . . . 20

vii

1.21 Model achitechture of a miniature hydrogen peroxide monopropellant

thruster, with a cross sectional view of the catalyst assembly [82] . . . . 22

1.22 Micro-bipropellant thruster from MIT [85] . . . . . . . . . . . . . . . . . 23

1.23 Schematic view of one solid propellant thruster [86] . . . . . . . . . . . . 24

1.24 Schematic of a conventional hybrid rocket motor [91] . . . . . . . . . . 25

1.25 (a) Vortex flow pancake hybrid model diagram [76] (b) Swirling of propel-

lant in a vortex flow [101] . . . . . . . . . . . . . . . . . . . . . . . . . . 30

2.1 Moles of hydrogen gas produced and the enthalpy of reaction against the

moles of reacted aluminium respectively . . . . . . . . . . . . . . . . . . 35

2.2 Selected energy density of some fuels [122] . . . . . . . . . . . . . . . . . . 37

2.3 Control volume with an attached nozzle: 𝐴𝑒 is the exit area of the nozzle,

𝐴𝑡 is the throat area, 𝑝𝑎 is the ambient pressure and 𝑝𝑒 is the exit pressure 41

2.4 Theoretical specific impulse performance against the expansion ratio of

the nozzle and the propellant mass flow rate . . . . . . . . . . . . . . . . . 47

2.5 ΔV performance versus the dry mass fraction for an 𝐼𝑠𝑝 of 118 𝑠 . . . . 48

2.6 Design model of the reaction chamber . . . . . . . . . . . . . . . . . . . 50

2.7 Swagelok cap and plug [143] adopted as nozzle . . . . . . . . . . . . . . . 51

3.1 Schematic of a pressure feed system . . . . . . . . . . . . . . . . . . . . 53

3.2 Feed system setup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

3.3 Circuit connection of the arduino and the solenoid valves . . . . . . . . 54

3.4 Thermocouples: (a) Insulated thermocouple and (b) Fine wire thermocouple 55

3.5 Pressure transducer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

3.6 DAQ measurement hardware and connections . . . . . . . . . . . . . . . 56

3.7 Block diagram of the LabVIEW program used to control and acquire data

from sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

3.8 𝑃𝑒𝑔𝑎𝑠𝑢𝑠 vacuum chamber . . . . . . . . . . . . . . . . . . . . . . . . . . 58

3.9 Schematic of the thrust stand . . . . . . . . . . . . . . . . . . . . . . . . 59

3.10 DC stepper motor from RS Components Limited . . . . . . . . . . . . . 60

3.11 Geometry Of Thrust Calibration and Thrust Stand Sepup with Stepper

Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

3.12 Responses of Thrust Calibration . . . . . . . . . . . . . . . . . . . . . . . 61

3.13 Schematic of the complete experimental setup . . . . . . . . . . . . . . . 63

viii

4.1 Experimental setup for sodium hydroxide concentration on aluminium-

water reaction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

4.2 Effect of sodium hydroxide molality on aluminium-water reaction . . . . 66

4.3 Schematic of the initial lab setup . . . . . . . . . . . . . . . . . . . . . . . 67

4.4 Changes in temperature and pressure in the reaction for the same mass

and concentration of oxidiser but different mass of fuel . . . . . . . . . . 68

4.5 Thrust reponses for different oxidiser concentration . . . . . . . . . . . . 70

4.6 Temperature reponses for different oxidiser concentration . . . . . . . . . 71

4.7 Average thrust performance response to oxidiser concentration . . . . . 72

4.8 One-shot thrust characterisation of the propulsion system on different

propellant ratios . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73

4.9 Temperature reponses for different propellant ratios . . . . . . . . . . . 74

4.10 Thruster characterisation at different propellant mass combinations . . . 75

4.11 Temperature reponses for different propellant mass combinations . . . . 76

4.12 Scaling of propellant ratio for more repeat cycles . . . . . . . . . . . . . . 77

4.13 Temperature reponses for different propellant mass combinations . . . . 78

4.14 Thrust level performance for different nozzle throat diameter . . . . . . 79

4.15 Temperature reponses for different nozzle throat diameters . . . . . . . 79

4.16 Reaction pattern of the thruster parameters . . . . . . . . . . . . . . . . . 81

4.17 Thrust, temperature and pressure responses of a one-shot experiment for

energy efficiency analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 82

4.18 Illustration of energy conversion efficiency . . . . . . . . . . . . . . . . . 83

4.19 𝑝 − ℎ diagram of water showing the enthalpy-pressure relation in the

reaction chamber. The 𝑝 − ℎ diagram was drawn from data obtained

from [156]. The blue line represents saturated liquid water while the

red line represents dry saturated steam. The dome covers water-steam

composition with decreasing water content from left to right. . . . . . . 85

4.20 Energy iterations for the percentage of water vapour in the system . . . . 87

4.21 Physical examination of propellant residue . . . . . . . . . . . . . . . . . 88

4.22 Microstructure view of two propellant residues . . . . . . . . . . . . . . 89

4.23 MicroRamam spectrum analysis of the propellant residues . . . . . . . . 89

4.24 Schematic layout of the propulsion system . . . . . . . . . . . . . . . . . 93

ix

4.25 CAD drawing of hybrid propuldion system for CubeSat applications

showing its dimensions in a 1𝑈 CubeSat . . . . . . . . . . . . . . . . . . 94

5.1 Uncertainty in repeated experimental data . . . . . . . . . . . . . . . . . 98

A.1 Reaction chamber . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 118

A.2 Nozzle part 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 119

A.3 Nozzle part 2 (Modified from swagelok [143]) . . . . . . . . . . . . . . . 120

A.4 Thrust attachment to the thrust stand . . . . . . . . . . . . . . . . . . . . 121

B.1 Aluminium wool data sheet . . . . . . . . . . . . . . . . . . . . . . . . . 123

B.2 K-type insulated thermocouple data sheet . . . . . . . . . . . . . . . . . 124

B.3 PFA needle valve data page 1 . . . . . . . . . . . . . . . . . . . . . . . . 125


B.5 PFA needle valve data page 3 . . . . . . . . . . . . . . . . . . . . . . . . . 127


B.7 DC stepper motor data sheet . . . . . . . . . . . . . . . . . . . . . . . . 129

B.8 Swagelok cap and plug data sheet . . . . . . . . . . . . . . . . . . . . . . 130

B.9 Solenoid valve description and data sheet 1 . . . . . . . . . . . . . . . . . 131

B.10 Solenoid valve description and data sheet 2 . . . . . . . . . . . . . . . . 132

B.11 Arduino Uno SMD Rev3 data sheet . . . . . . . . . . . . . . . . . . . . 133

B.12 Laser displacement sensor (optoNCDT 1700-50) data page 1 . . . . . . . 134

B.13 Laser displacement sensor (optoNCDT 1700-50) data page 2 . . . . . . . 135

x

List of Tables

1.1 Performance characteristics of propulsion systems flown on board CubeSats 5

1.2 Classification of satellites showing 1𝑈 CubeSat limited resources . . . . 6

1.3 Performance comparison of micropropulsions as requirements for CubeSat

propulsion system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

1.4 Propulsion requirements for nanosatellites for several missions [92] . . . . 27

1.5 Chemical propulsion trade-off for CubeSat applications . . . . . . . . . . 28

1.6 Thermochemical analysis of propellant combinations . . . . . . . . . . . . 31

2.1 Thermodynamic properties of the propellants . . . . . . . . . . . . . . . 36

2.2 Heat capacity coefficients . . . . . . . . . . . . . . . . . . . . . . . . . . 40

2.3 Hybrid propulsion system design parameters and performance . . . . . . 48

4.1 Experiment data for sodium hydroxide molality . . . . . . . . . . . . . . 65

4.2 Temperature and pressure rise in vacuum condition . . . . . . . . . . . . . 67

4.3 Data for effect of oxidiser concentration on thrust level . . . . . . . . . . 69

4.4 Data for fuel/oxidiser effect on one-shot experiment . . . . . . . . . . . 72

4.5 Experimental data for variation in propellant mass at constant ratio . . 75

4.6 Experimental data for scaling effect and repeat cycles . . . . . . . . . . . 77

4.7 Experimental data on the effect of different nozzle throat diameter . . . 79

4.8 Experimental data for the chemical reaction model of the thruster . . . 80

4.9 Data for a one-shot experiment . . . . . . . . . . . . . . . . . . . . . . . 82

4.10 Thermodynamic properties of the propellants . . . . . . . . . . . . . . . 84

4.11 Table of comparison between theory and experimental data . . . . . . . . 91

4.12 Mass estimation of the hybrid propulsion system . . . . . . . . . . . . . 95

xi

Nomenclature

Arcjet

Bipropellant Systems

Hybrid Propulsion System

Monopropellant Systems

Nanosatellites

Resistojet

Solid Rocket Motors

Vacuum Arc Thruster

Δℎ∘𝑓(𝑝𝑟𝑡𝑠) Specific standard heat of formation of the products, 𝑘𝐽/𝑘𝑔

Δℎ∘𝑓(𝑟𝑐𝑡𝑡𝑠) Specific standard heat of formation of the reactants, 𝑘𝐽/𝑘𝑔

Δℎ∘𝑟𝑥𝑛 Specific enthalpy of formation, 𝑘𝐽/𝑘𝑔

ΔV Net velocity change to the spacecraft,𝑚/𝑠

𝑒 Outlet mass flow rate, 𝑘𝑔/𝑠

𝑖 Inlet mass flow rate, 𝑘𝑔/𝑠

𝑐𝑣 Net rate of energy transfer by heat across the boundary of the

control volume, 𝐽/𝑠

𝑐𝑣 Net rate of energy transfer by work across the boundary of the

control volume, 𝐽/𝑠

𝛾 Ratio of specific heat capacities

𝜆 Thrust efficiency,%

𝑎 𝑏 𝑐 Heat capacity coefficients

𝑎𝑜 Acoustic velocity,𝑚/𝑠

𝐴𝑡 Nozzle throat area,𝑚2

𝐴𝐹𝑅𝐿 Air Force Research Laboratory

𝐴𝑇𝐼 Advanced Technology Institute

xii

𝑐 Effective exhaust velocity,𝑚/𝑠

𝑐* Characteristic velocity,𝑚/𝑠

𝑐𝐹 Coefficient of thrust

𝑐𝑝 Temperature-dependent heat capacity at constant pressure,𝐽

𝐶𝐴𝑁𝑋 − 2 Canadian Advanced Nanospace Experiment-2

𝐶𝐸𝐴 Chemical Equilibrium with Applications

𝐶𝐺𝐺 Cold Gas Generator

𝐶𝑀𝑂𝑆 Complementary Metal Oxide Semiconductor

𝐷𝐴𝑄 Data acquisition

𝑒 Nozzle area expansion ratio

𝐸𝑐𝑣 Energy of the control volume,𝐽

𝑒𝐿𝐼𝑆𝐴 Evolved Laser Interferometer Space Antenna

𝐸𝑆𝐴 European Space Agency

𝐹 Steady thrust force,𝑁

𝐹ℎ Holintal force,𝑁

𝐹𝐸𝐸𝑃 Field Emission Electric

𝐹𝑀𝑀𝑅 Free-Molecule Micro-Resistojet

𝑔 Acceleration due to gravity at the surface of the earth,𝑚/𝑠2

ℎ𝑒 Total specific enthalpy of outlet from the control volume, 𝑘𝐽/𝑘𝑔

ℎ𝑖 Total specific enthalpy of inlet to the control volume, 𝑘𝐽/𝑘𝑔

ℎ𝑐ℎ𝑒𝑚 Specific enthalpy of chemical reaction, 𝑘𝐽/𝑘𝑔

𝐻𝑇𝑃 High Test Peroxide

𝐻𝑇𝑃𝐵 Hydroxyl-Terminated Polybutadiene

𝐼𝑠𝑝 Specific impulse, 𝑠

𝐼𝑡𝑜𝑡 Total impulse,𝑁𝑠

𝐼𝑂𝑁 Illinois Observation Nanosatellite

𝐽𝑃𝐿 Jet Propulsion Laboratory

𝑙 Length of pendulum thread suspending the calibration mass,𝑚

𝐿𝑂 Liquid Oxygen

𝑀 Mach number

𝑚𝑐 Calibration mass, 𝑘𝑔

𝑚𝑓V Final mass of the satellite after the ejection of propellant, 𝑘𝑔

𝑚𝑖V Initial mass of the satellite including the propellant, 𝑘𝑔

xiii

𝑀𝑚 Gas molecular mass, 𝑘𝑔/𝑘𝑚𝑜𝑙

𝑀𝐶𝐷𝑇 Microcavity Discharge Thruster

𝑀𝐸𝑀𝑆 Micro-Electro-Mechanical Systems

𝑀𝑀𝐻 Mono-methyl-hydrazine

𝑁2𝐻4 Hydrazine

𝑁𝐴𝑁𝑂𝑃𝑆 Nano Propulsion System

𝑁𝐴𝑆𝐴 National Aeronautics and Space Administration

𝑁𝑇𝑂 Nitrogen-tetroxide

𝑝𝑜 Stagnation pressure, 𝑘𝑃𝑎

𝑃𝐴𝐶 Primex Aerospace Company

𝑃𝐸 Polyethylene

𝑃𝐸𝐸𝐾 Polyether-Ether-Ketone

𝑃𝑀𝑀𝐴 Poly-Methyl Methacrylate

𝑃𝑃𝑇 Pulsed Plasma Thruster

𝑃𝑃𝑈 Power Processing Unit

𝑃𝑇𝐹𝐸 Polytetrafluoroethylene

𝑝𝑣 Flow work,𝐽

𝑅 Universal gas constant, 𝐽/𝑘𝑚𝑜𝑙.𝐾

𝑅𝑜 Specific gas constant,𝐽/𝑘𝑔𝐾

𝑠 Horizontal displacement distance of mass from thrust stand,𝑚

𝑆𝐸𝐸 Secondary Electron Emission

𝑆𝐹6 Sulphur Hexafluoride

𝑆𝑃𝑇 Solid propellant thruster

𝑆𝑃𝑇 Stationary Plasma Thruster

𝑆𝑇𝑅𝑎𝑁𝐷 − 1 Surrey Training, Research and Nanosatellite Demonstrator-1

𝑇 Temperature at the point of interest the stagnation streamline,∘C

𝑡 Time, 𝑠

𝑇𝑜 Stagnation temperature, ∘C

𝑇𝐴𝐿 Thruster with Anode Layer

𝑢 Temperature-dependent internal energy, 𝐽

𝑉𝑒 Exhaust velocity,𝑚/𝑠

𝑣𝑒 Gas exit velocity,𝑚/𝑠

𝑣𝑖 Inlet velocity of the flow,𝑚/𝑠

xiv

𝑉 𝐿𝑇 Vaporizing Liquid Thruster

𝑋𝑅𝐷 X-ray power diffraction

𝑧𝑒 Vertical measurement of outlet from the control volume,𝑚

𝑧𝑖 Vertical measurement of inlet to the control volume,𝑚

H2O2 Hydrogen peroxide

LV Launch Vehicle

N2O Nitrogen oxide

xv

Chapter 1

Introduction

1.1 CubeSat

First standardised in 1999 by Professor Jordi Puig-Suari at California Polytechnic State

University and Professor Bob Twiggs at Stanford University [1], CubeSat is normally

represented in different configurations as 1𝑈 , 2𝑈 , 3𝑈 or more, with each 𝑈 being a 10 𝑐𝑚

cube size with a volume of 1𝐿 and 1.330 𝑘𝑔 mass for a 1𝑈 CubeSat [2, 3], and has an

average available power of 1.6𝑊 (about 100𝑊/𝑚2 from the body mounted solar panels)

for all the subsystems [4, 1]. An example of a CubeSat is shown in Figure 1.1.

Figure 1.1: 1U cube satellite [5]

The intention of the early nanosatellites, and in particular CubeSats, was primarily for

students of higher learning to have a hands - on experience in designing, developing,

testing and operating satellite systems [6]. These satellites were built and launched

into space without propulsion system, and their orbit naturally decayed and deorbit

1

1.1. CubeSat

into the atmosphere, and this has restricted the altitude of the nanosatellites to less

than 400 km in order to deorbit within the regulated 25 years without creating space

debris [7]. In some cases their attitude control was performed using magnetic torquers

and momentum wheels [8]. CubeSats are now becoming increasingly popular among

universities, governmental and non-governmental organisations such as European Space

Agency (ESA) [9] and National Aeronautics and Space Administration (NASA) [10], and

not just for university teaching tools, but for the purposes of earth observation, scientific

and technology demonstrations, surveillance, global positioning system navigation and

communication [11, 12]. This is due to the design, build and launch costs of these

satellites, recent changes in government policies and rapid advances in decreasing satellite

electronics size with increased capability at very low power consumption [13, 14]. The

introduction of these satellites could improve satellite reliability and performance as

the functionality of a traditional satellite will be undertaken by several microsatellites

woking in parallel [15]. A good example of such a constellation flight program is the

QB50 [16]. Figure 1.2 is an example of a constellation flight.

Figure 1.2: Concept of constellation flight [17]

Apart from redistributing the tasks of bigger satellites, nanosatellites could take more

complex missions in higher altitude with limited cost and planning time, leading to

rapid developmental timetables by using commercial off the shelf (COTS) technology

2

1.1. CubeSat

[18, 15]. Limitations to this space technology development include: 1) miniaturisation of

conventional propulsion systems that would enable the spacecraft to take complex missions

in higher altitude (> 400 𝑘𝑚). This is because scaling well understood conventional

propulsion systems to the constraint size, mass, power and energy limitations of the

nanosatellites while still retaining their operation advantages and performances is difficult

and complex [12, 15, 19]; 2) the second limitation is the requirement to de-orbit the

satellite after the end of life operation, and within the 25 years regulated period to avoid

space debris.

1.1.1 CubeSats Flown with Propulsion Systems

Growing interest in CubeSats has necessitated the inclusion of micropropulsion systems

to expand their area of application. In the last decade for instance, there are only three

CubeSats that have been successfully flown with propulsion systems on board. These

include:

1. 3𝑈 CanX-2 CubeSat which was designed by the University of Toronto in 2008 and

launched on the Indian Polar Satellite Launch Vehicle. The mission objective of

the CubeSat was to demonstrate several enabling technologies for precise formation

flight [20] and in space inspection on the smallest platform possible and to perform

risk mitigation assessment for the critical components for CanX-4 and -5 missions

[21]. It incorporated a liquid-fuelled cold gas Nano Propulsion System (NANOPS),

see Figure 1.3, that used sulphur hexafluoride (SF6) as the propellant [22, 20] and

the propulsion system was estimated to deliver about 35𝑚/𝑠 ΔV to the CubeSat.

Other performances of the propulsion system are shown Table 1.1.

Figure 1.3: CanX-2 NANOPS system [20]

3

1.1. CubeSat

2. The second CubeSat flown with propulsion system on board is the STRaND-1

designed and developed by SSTL and SSC to demonstrate new technology in

space using a smart phone as the primary payload. The CubeSat was launched on

25th February, 2013 and it represents the first smartphone on a nanosatellite in

space [23, 24]. The 3𝑈 CubeSat had two propulsion systems on board: a butane

resistojet which was designed to provide 2𝑚/𝑠 ΔV to the satellite and a pulsed

plasma thruster (PPT) of 1340 𝑠 specific impulse with a total of 2.7𝑚/𝑠 ΔV [25],

see Figure 1.4. Other performances of the STRaND-1 propulsion system are listed

in Table 1.1.

(a) (b)

Figure 1.4: STRaND-1 propulsion systems: (a) Butane resistojet (b) Pulsed plasmathruster [25]

3. The third CubeSat is the Delfi-n3Xt, a 3𝑈 Dutch picosatellite that is operated

by the Delft University of Technology Netherlands [26]. The CubeSat, designed

to demonstrate propulsion and communication systems for future missions was

launched with a Dnepr launch vehicle in Russia on 21st November, 2013 [27]. The

micropropulsion system on board the CubeSat is a cold gas generator (CGG)

that stores nitrogen in a solidified form but turns to gas on operation [28, 29],

and requires about 11𝑊 of power for operation, see Figure 1.5. The CubeSat is

expected to perform orbital change by the operation of the micropropulsion system.

The performance characteristics of these propulsion systems are shown in Table

1.1.

4

1.1. CubeSat

Figure 1.5: Delfi-n3Xt cold gas generator thruster components [29]

Table 1.1: Performance characteristics of propulsion systems flown on board CubeSats

Cubesat Propulsion Propellant Mass, Specific Thrust, Total ΔV, Power,

type kg impulse, mN impulse, m/s W

s Ns

CanX-2 Cold gas 𝑆𝐹6 0.50 50-100 50-100 - <35 -

STRaND-1 Resistojet; Butane 0.50 90; 100; - 4.7 ≈8.6

PPT Capacitor 1340 0.0009

Electrodes

Delfi-n3Xt Cold gas Nitrogen 0.12 ≈30 6-100 - - 11.018

Generator

About 60% of the flown propulsion systems in all the 3𝑈 CubeSata is cold gas thrusters,

which is drawn from the low cost and simplicity of the systems. However, continuous

leakages of the propellants along the mission stage, resulting in performance reduction,

hinders the deployment of these systems to de-orbit the nanosatellites in a mission

above the low earth orbit. The resistojet and the cold gas generator power requirements

are essentially beyond the available power onboard a 1𝑈 CubeSat, and the propulsion

systems also have low ΔV performances limiting the missions that the nanosatellites can

undertake.

There is still a need for nanosatellites propulsion system that is safe and simple, cheap,

non-toxic, requires power that is less than it is available on a 1𝑈 CubeSat, storable,

repeatable, operating within the temperature range of nanosatellite, and yet can de-

liver high ΔV for orbital maintenance. This is what this research addresses. However,

nanosatellittes likes a 1𝑈 CubeSat is constraint in mass, volume and power as shown in

Table 1.2 requires propulsion system that is within the budget and constraint resources

on the CubeSat and whose propellants are readily available and safe. This will enables

5

1.2. Overview of Micropropulsion Systems

the nanosatellite to contend with bigger satellites in mission.

Table 1.2: Classification of satellites showing 1𝑈 CubeSat limited resources

Group name Mass (kg)

Large satellite >1000

Medium satellite 500 to 1000

Minisatellite 100 to 500

Microsatellite 10 to 100 Mass (kg) Volume (m3) Power (W)

Nanosatellite 1 to 10 CubSat 1.33 0.001 1.6

Picosatellite 0.1 to 1

Femtosatellite <1

1.2 Overview of Micropropulsion Systems

The major constraints to the deployment of available chemical and electric micropropul-

sion systems on nanosatellites missions are safety issues, cost, high power/energy demand

and highly complex subsystem. Safety in term of hazardous propellants - the propellants

used in monopropellant and bipropellant systems are highly toxic, which increases the

overall cost of a CubeSat mission because of high handling cost. These propellants include

high test peroxide, hydrazine, mono-methyl-hydrazine (MMH) and nitrogen-tetroxide

(NTO); and safety in term of high pressure cold gas system - high pressuure cold gas

systems cannot be flown with other primary paylaods according to NASA and Johnson

Space Center [30]. Also, the development of reliable power processing unit for electric

micropropulsion systems is complex and expensive, aside the fact that these systems are

power dependant and energy limited for nanosatellites. For example, propulsion systems

by BUSEK for nanosatellites have input power in the range of 3.5𝑊 to 15𝑊 [31]. Cur-

rent development in micropropulsion systems, their operations, their performances and

their suitability for low cost nanosatellite missions are highlighted in this section.

1.2.1 Electric Micropropulsion Systems

All the electric propulsion systems operate by adding energy to the working fluid from

an electric source to ionise and/or accelerate the propellant to provide thrust [32]. The

6


energy is processed by a subunit, which is complex to design. The generated thrust is

related to the input power by 𝐹 = 𝑃𝑖𝑛2𝜂𝑣𝑒, where 𝜂 is the power conversion efficiency

and 𝑣𝑒 is the propellant exit velocity. The thrust generated by this system is small due

to limited electric energy on-board the microsarellites. However, the thrusting time is

long with high propellant utilisation efficient with fine impulse. Electric propulsion is

further classified into electrothermal, electrostatic and electromagnetic according to the

acceleration of the propellant out of the system.

1.2.1.1 Resistojets

Resistojets are examples of electrothermal propulsion systems in which propellant is

heated through direct ohmic heating by passing it over a very hot metal element to elevate

the propellant temperature before passing it through an exhaust nozzle to generate thrust.

Resistojets use different working fluids as propellants ranging from water (𝐻2𝑂) [33],

ammonia (𝑁𝐻3), high test peroxide (𝐻𝑇𝑃 ) and hydrazine (𝑁2𝐻4) which also determine

their specific impulse and thrust performance levels [5]. Factors affecting the choice

of these propellants also include cost of propellants, ease of catalytic decomposition

and environmental and health concerns [12] For instance, a 1𝑘𝑔 of ammonia cost about

$0.31 and a 1 𝑘𝑔 of HTP cost about $0.17, whereas a 1 𝑘𝑔 of hydrazine cost about $17

[34, 35, 36]. In recent past, research has focused on how to miniaturise the technology

using micro-electro-mechanical systems (MEMS)-fabrication techniques for their potential

applications in nanosatellites. Among such efforts is a vaporizing liquid thruster (VLT)

developed by Mukerjee 𝑒𝑡 𝑎𝑙 [37] and shown in Figure 1.6.

(a) Schematic view of the microthruster[38] (b) Complete mi-crothruster [37]

Figure 1.6: Schematic and complete views of vaporizing liquid thruster

It is operated by injecting the liquid propellant (water or hydrazine) into a micro-

7


machined micro-chamber containing silicon heaters where it is vaporises and passes

through a micro-silicon nozzle to produce thrust. They have recorded an initial thrust

performance of 0.15𝑚𝑁 to 0.46𝑚𝑁 at an operating power of 5𝑊 to 10.8𝑊 with a

propellant input flow rate of about 0.09 𝑐𝑐/𝑠. A similar design was investigated by

Mueller 𝑒𝑡 𝑎𝑙 [39] at the NASA’s Jet Propulsion Laboratory (JPL). They used water as

the working fluid but at a heating power of 2𝑊 . Even at this power level, the thrust

value ranges between 50-280𝜇𝑁 with a thrust/power ratio of 200𝜇𝑁/𝑊 and at a specific

impulse of about 100 𝑠 while operating at low feed pressure.

Complementary metal oxide semiconductor (CMOS) resistojet from Janson at the

Aerospace Corporation, California [40] is another effort of making electro-thermal system

through batch-fabrication of MEMS. The heating element is provided by a polysilicon

layer sandwiched between 2 patterned passivation layers, which normally acts as gate

structure in a CMOS transistor [41]. After many iterations of development, they devel-

oped a 3-Watt CMOS microresistojet that incorporates a flow sensor and low resistance

power transistor as shown in Figure 1.7.

Resistojet 3

Flow RateMonitor

PowerTransistor

Poly Heater

Inlet

Plenum

Nozzle

TOPHALF:

BOTTOMHALF:

Pads

Figure 1.7: 3-Watt CMOS resistojet on a 2-micron TinyChip die layout [40]

Performance for the CMOS resistojet in literature was reported by Maurya 𝑒𝑡 𝑎𝑙 [42] where

they demonstrated thrust range of 5-120𝜇𝑁 at a heating power of 1-2.4𝑊 . Free-Molecule

Micro-Resistojet (FMMR) is another MEMS-based resistojet developed by Ketsdever 𝑒𝑡

𝑎𝑙 [43] at the Air Force Research Laboratory, California. The principle of operation is

similar to that of CMOS resistojet but has offered a higher thrust performance of near

0.25𝑚𝑁 at a specific impulse of almost 45 𝑠 when operated at a stagnation temperature

8


of 600𝐾 using argon propellant. The major advantage of the MEMS-based resistojet

is their small size and weight, scalability ability with a very precise thrust impulse.

However, the power processing unit is complex and expensive for nanosatellites.

1.2.1.2 Arcjets

Arcjet thrusters are also electro-thermal propulsion systems but they use an arc discharge

through the supply of high voltage across an anode and a cathode to ionise the propellant.

This allows the passage of DC current through the ionised gas that heats up the propellant

into directed flow by increasing the propellant and transfers this energy into directed

flow by increasing the propellant kinetic energy. The superheated gas is directed through

a nozzle to create thrust. The most common propellants are hydrogen for ground testing,

and ammonia and hydrazine for in flight applications [44]. Figure 1.8 is a schematic

drawing of a laboratory model of an arcjet thruster.

Figure 1.8: Schematic of an arcjet thruster [44]

Horisawa 𝑒𝑡 𝑎𝑙 [45] have used laser machined technology to design a micro-arcjet for

microsatellites applications with extensive work on the thruster micro-nozzle manufacture.

Their test result showed a thrust level of 1.2𝑚𝑁 and specific impulse of 147 𝑠 with a

thrust efficiency of 7% at an input power of 6𝑊 . The major drawbacks to the micro-arcjet

are heat transfer and energy burden issue on nanosatellites [46].

1.2.1.3 Microcavity Discharge Thruster

Microcavity discharge thruster (MCDT) is another electro thermal propulsion concept by

the University of Illinois. A gaseous propellant is supplied through a 70-130𝜇𝑚 diameter

cavity that is created though the bonding of a two insulated Al/Al2O3 electrodes that are

9


powered by a 50-150 𝑘𝐻𝑧 and 400-1200𝑉 𝐴𝐶 source [47]. Figure 1.9 shows the schematic

of the insulated electrodes and the microcavity. The cavity holds the discharge plasma

and the plasma is pressurised to a pressure of about 1 𝑎𝑡𝑚 at a temperature reaching

1500𝐾, thereby expanding the gas through a micro-nozzle to produce thrust. Unlike the

arcjet, it uses an alternating voltage to create an alternating electric field in the cavity

to partially ionise the gas. Scalability of the device has been performed by using the

thrusters in an array, with over 1 million arrays of cavities.

Al2O3

Al

Figure 1.9: Schematic of an insulated electrodes of a microcavity thruster [47]

Initial thrust performance through experimental measurements with a 0.25𝑊 per cavity

has been demonstrated using neon as propellant, and using water vapour and nitrogen

gas to enhance power utilisation [4]. So far, they have recorded 0.6-2.7𝑚𝑁 of thrust at

a pressure range of 120-240 𝑘𝑃𝑎 and propellant flow rate of 0.99-5.22𝑚𝑔/𝑠 for a 4 cavity

array with 120𝜇𝑚 and 210𝜇𝑚 throat and exit plane diameters respectively. The power

processing unit for such a technology remains a challenge and is the focus of the ongoing

work.

1.2.1.4 Ion Engines

Ion engines are electrostatic systems where ions are extracted from low-pressure plasma

through an electrostatic grid to a high exit velocity of about 30,000𝑚/𝑠 [38]. The

generation of the plasma from the propellant could be through microwave heating, hollow

cathode electron emission or through radio frequency plasma excitation, and are classified

as DC electron bombardment or Kaufman-type thrusters and RF ion engines [4, 15].

Several studies have been done to miniaturise the thruster as reported by Mueller 𝑒𝑡 𝑎𝑙

[4] but as the ionization chamber becomes smaller, the ion production decreases resulting

in reduced thrust efficiency [15]. One of the recent developments in ion engines for

10


microsatellites and on lower power scale is the work of Taunay 𝑒𝑡 𝑎𝑙 [48] at Pennsylvania

State University, USA. They conducted computational and experimental studies on both

the radio frequency and microwave ion thrusters using argon and xenon propellants.

Figure 1.10 shows the schematic diagrams of a 3𝐷 view and a longitudinal cross-section

of their thruster. They achieved a thrust of 59𝜇𝑁 for the radio frequency thruster at

a specific impulse of 5,480 𝑠 operating within 13𝑊 of power input with a mass flow

rates of between 0.02 𝑠𝑐𝑐𝑚 and 0.1 𝑠𝑐𝑐𝑚. The microwave ion thruster was run on similar

mass flow rates but it consumed a lower power of 8𝑊 . They recorded higher thrust

performance of 217𝜇𝑁 at a specific impulse of 10,700 𝑠, with higher total efficiency of

66.3%.

Figure 1.10: Schematic diagrams of a 3D view and a longitudinal cross-section view ofthe ion thruster developed at Pennsylvania State University [48]

Ion engines require high operating energy that the nanosatellites cannot provide.

1.2.1.5 Hall Thrusters

The hall thrusters operate by the attraction of electrons that are produced from an

external hollow cathode toward the main annular chamber by a metal anode, see Figure

1.11. The electrons are then subjected to a radial magnetic field that is established

by electromagnets in the thruster. Due to an 𝐸 × 𝐵 effect on the electrons by both

the electric and magnetic fields, the electrons then travel in a circular azimuthal Hall

drift resulting in the ionization of an inert propellant that is fed from the anode to the

thruster chamber. The ions, which have a higher mass to charge ratio than the electrons,

are then accelerated out of the thruster at a very high velocity by the potential difference

11


across the magnetic field to produce thrust. There are two main types of hall thrusters:

stationary plasma thruster (SPT) and thruster with anode layer (TAL), with their major

difference in the construction materials of their channels [15]. While the SPT is made of

boron nitride walls, the TAL is made of stainless steel with a resultant effect in secondary

electron emission (SEE) [49].

Figure 1.11: Schematic diagram of an SPT Hall thruster, showing the electrodes and theradial magnetic field [49]

A novel cylindrical-type and lower-power miniaturised Hall thruster operating at an

input power of about 10𝑊 is the one developed by Ikeda 𝑒𝑡 𝑎𝑙 [50] at the Osaka Institute

of Technology, Japan. Tagged TCHT-4, as shown in Figure 1.12, the thruster was born

after some iteration of previous series of TCHT to improve the power utilisation. They

indicated thrust performance of up to 7.3𝑚𝑁 and a specific impulse of 940 𝑠 at an input

power of 10𝑊 , power which is high above the typical power rating of nanosatellite.

12


W [2][3]. Detailed effects of magnetic fieldcylindrical Hall thrusters are unknown,

ch important to improve thrust performance.tigated the effects with the cylindrical Hall

d TCHT series in Osaka Institute ofdischarge chamber consists of only a circular

art with no coaxial parts. Although cylindricalde by Raitses and Smirnov have short coaxialpart was excluded from TCHT-series. Byradial magnetic field at the downstream

ruster TCHT-3B achieved higher thrustn TCHT-3A did at low power level because ofll losses[4]-[10]. However, when the position

( )

Figure 1. Cross-sectional view of TCHT-4.

Figure 2. Photo of TCHT-4.

Ceramic Wall

PermanentMagnet

Anode

(mm)

Hollow Cathode

Propellant

CoilSm-Co MagnetCopperBoron NitrideAluminumIron

27

35

0

0

Figure 1.12: Schematic diagram and photo shot of a low-power miniaturised Hall thruster(TCHT-4) [50]

1.2.1.6 Micro Pulse Plasma Thrusters

Micro pulse plasma thrusters (𝜇PPTs) are electromagnetic thrusters whose principle of

operation is similar to the conventional PPTs, where a bar of Teflon propellant is placed

between two electrodes with a spring that pushes the propellant for consumption. A

capacitor is charged to provide the required power that ionises fraction of the propellant

into plasma when ignited by a spark plug. The plasma is then accelerated to a very high

velocity to produce thrust [51, 38]. The 𝜇PPT uses a coaxial fuel rod and eliminates the

use of the igniter spark plug, trigger electronics, propellant housing structure, propellant

spring and with half of the power processing unit [52].

Spanjers 𝑒𝑡 𝑎𝑙 [52] has developed two types of Air Force Research Laboratory (AFRL)

𝜇PPT, shown in Figure 1.13(a), that use surface discharge across a Teflon propellant

in two or three overlapping conducting electrodes that are self-triggered. AFRL 𝜇PPT

thrust performance is an average of 2-30𝜇𝑁 when operated at an input power of 1-

20𝑊 . This gives a total impulse of 2𝜇𝑁𝑠 per shot at a discharge energy of about 1 𝐽

per pulse [4, 52]. Also in 2005, a 𝜇PPT termed Dawgstar thruster was developed by

Primex Aerospace in partnership with University of Washington (Cornell University)

for a university nanosatellite project. This design was similar to a conventional PPT

but miniaturised to provide orbital maintenance for the satellite. The recorded thrust

level performance for the thruster is 60-275𝜇𝑁 with a specific impulse of 266𝑠 at input

13


power of 15.6-36𝑊 depending on charging rate and thrust frequencies [53]. Another

𝜇PPT, Figure 1.13(b), is being investigated in the UK by a collaboration between the

University of Southampton, Mars Space Ltd and Clyde Space Ltd. The main objective

of the project is to double the life span of a CubeSat when launched into LEO orbits of

altitude 600-650 𝑘𝑚 when included in its design [54]. Though their system has shown

a satisfactory results with a specific impulse of 640 𝑠 and a thruster mass of 500 𝑔, the

operating power for a single thrust unit is 10𝑊 [55].

(a) Schematic of an AFRL 3-electrode𝜇PPT concept [52]

(b) Assembled breech-fed 𝜇PPT[55]

Figure 1.13: 𝜇PPT for concepts for microsatellites

The major advantages of 𝜇PPT are their simplicity in design, high reliability, and

durability, but create electromagnetic interference for other payloads and high voltage

operation have so far limited their application [56].

1.2.1.7 Micro Laser Ablation Thruster

(a) Schematic diagram of a micro laserablation thruster

(b) Micro laser thruster testbed

Figure 1.14: Micro laser ablation thruster concept [57]

14


The micro laser ablation thruster, also known as micro laser plasma thruster, uses laser

diode technology [57] to produce thrust from an ablation target. Figure 1.14 shows the

operation principle of a micro laser ablation thruster and the thruster testbed. Lenses

are used to focus the diodes laser beams on the ablation target (a two-layer fuel tape)

with a transparent supporting layer upon which the laser light passes to produce a very

small jets of plasma that results in thrust by igniting an absorbing fuel layer [ 58]. The

operation of the motor provides a successive layer of the tape for the laser light for

ablation. Performance characteristic of the device indicated a thrust of 680𝜇𝑁 at a

specific impulse of about 400 𝑠 when operating with an optical power of 2.1𝑊 and 15𝑊

peak power of laser diode at a tape lifetime of 140ℎ [57]. It’s potential application is

in the area of precise attitude control in constellation for high accuracy interferometer

mission like the Evolved Laser Interferometer Space Antenna (eLISA) [59]. However, it

requires a complex and high input power for its operation.

1.2.1.8 Vacuum Arc Thruster

The vacuum arc thruster, developed by Alameda Applied Science Corp.(AASC) [56], is

an ablative pulse propulsion [4] system where a high voltage potential is applied across

two metal or non-metal electrodes in a vacuum. The applied potential creates an erosion

of the electrode (cathode) and ejects plasma at high velocity into an inter-electrode gap

[60]. The accelerated plasma is then ejected out of the vacuum to generate thrust. The

latest modification to the thruster by Keidar 𝑒𝑡 𝑎𝑙 [61] is the use of magnetic field across

the electrodes, see Figure 1.15, to enhance the power efficiency and increase the specific

impulse. The new thruster is known as magnetically enhanced vacuum arc thruster

(MVAT), and their performance has shown a 50% increase in power efficiency and 30%

increase in specific impulse [62].

CoreIsolator

Anode

Magnetic CoilSpring Cathode

Figure 1.15: Schematic diagram of a magnetically enhanced vacuum arc thruster [62]

15


Four micro vacuum arc thrusters were developed for the University of Illinois 2-cube

CubeSat-Illinois Observation Nanosatellite (ION), using a 150 𝑔 and 12-24𝑉 power

processing unit (PPU) that was designed and built by AASC [63]. The specific impulse

of the VAT was 3000𝑠 and its thrust to power ratio was about 10𝜇𝑁/𝑊 with an input

power ranging from 1-100𝑊 , depending on the required thrust. The satellite, which

was lost due to the failure of the launch vehicle in 2006, could not get to test the 2-axis

control and orbit translation abilities of the microVATs on board. The low thrust to

power ratio of VAT put a high power demand burden on nanosatellites for operations

involving large ΔV, with a limit on the thrust performance, even with the advantage of

high specific impulse.

1.2.1.9 Field Emission Electric Propulsion

Field emission electric Propulsion (FEEP) uses an electrospray technique [64, 15] where

the ions of heated liquid metal propellant (like indium) are extracted and accelerated

to produce thrust by the application of a high electric field using an emitter and an

extractor. Equilibrium between the applied electric field and the surface tension of the

liquid causes a Taylor cone at the tip of the emitter resulting in a protruding tip as

shown in Figure 1.16. A neutraliser is provided at the exit of the ejected ions to prevent

unbalance of electrical charges around the spacecraft. Both ionization and acceleration of

the ions are done by the same electric field and the propellant is not pressure fed except

by capillary force, making the propulsion system scalable for nanossatellites [4].

ions+

Emitter section

Slit or needle shape

Ve Va

Ibeam

Accelerator

Propellant tank

+ - e-

Neutralizer

Figure 1.16: Schematic diagram of a field emission electric propulsion [65]

A close counterpart of the FEEP is the colloid thruster. Unlike the emission of ions in

the FEEP, charged liquid droplets of the liquid propellant (doped glycerol) are caused to

16


break away due to the strong electric field across the electrodes [15]. Recent development

on the electrospray technology for a low power FEEP microthruster is the FT-150 FEEP

designed for LISA Pathfinder mission. The collaboration is between Astrium Space

Transportation in France and Austrian Research Centre and the thruster was to provide

fine positioning and attitude control on 𝜇N thrust range. The last iteration of the project

at ALTA SpA, Italy showed a thrust performance of 0.1-150 𝜇𝑁 and a specific impulse

range of 3000-4500 𝑠 at an input power of 6𝑊 [66]. Though the FEEP system is boastful

of high specific impulse, the thrust to power ratio is very low and it requires high energy

for orbital maintenance.

1.2.2 Chemical Micropropulsion Systems

In chemical propulsion systems, thrust is generated through thermodynamics using the

stored chemical energy in the propellants, and accelerating the ejected stream of gaseous

products through a converging and diverging nozzle to produce thrust. The generated

thrust, 𝐹 , is proportional to the product of the propellant mass flow rate, and the

exit velocity, 𝑣𝑒. That is, 𝐹 = 𝑝𝑣𝑒, and depending on the enthalpy and the pressure of

the chemical reaction, the thrust value can be moderate to high and occurring within a

short time. Chemical propulsion systems have flight heritage for attitude control and

orbital raising involving low to moderate ΔV requirements with a thrust-to-weight ratio

of 0.1− 0.3 [12] especially suitable for rapid orbital manoeuvres for traditional satellites.

However, their applications on nanosatellites missions have performance, safety, thermal

control and scaling concern issues that will be highlighted in this section. They are

classified into cold gas and hot gas propulsion systems based on the exhaust gas from the

nozzle. The propellant could be gaseous, liquid, solid or a combination of these.

1.2.2.1 Cold Gas Thruster

In a cold gas system, gas from a high-pressure tank, vaporised liquid or solidified gas is

vented through a valve and nozzle to produce thrust. Figure 1.17 shows the schematic

description of a cold gas thruster with a detailed view of the component parts.

17


Pressure gauge Gas Tank

PressureRegulator Stage 1

Fill/Drain Valve

Gas Filter

Relief Valve

PressureRegulator Stage 2

Accumulator Tank

Solenoid Valve

Nozzle

Figure 1.17: Schematic view of cold gas thruster [38]

The system is characterised by low power demand mainly for valve opening/closing

operations, simplicity, cleanliness, robustness and safety, though with low thrust and

low specific impulse performance ranging from 30 𝑠 to 100 𝑠 [67]. It is mainly used for

attitude control that requires small ΔV applications. Popular among cold gas thrusters

manufacturers is Moog Inc. who has been developing several cold gas thrusters and

their components for small satellites for the past over 15 years [68]. Due to the power

requirement for their valve operations, they remain unsuitable for nanosatellites [4].

Aerojet is also developing a CubeSat propulsion system from their Modular Propulsion

Systems product line, tagged 𝑀𝑃𝑆 − 110 [69] for CubeSat applications.

18


Figure 1.18: Schematic diagram of MPS-110 cold gas thruster developing by Aerojet [69]

The scalable thruster, shown in Figure 1.18, is to provide primary propulsion requirements

that involve minimal ΔV during constellation deployment, orbit maintenance and end of

life de-orbiting of the CubeSats [70]. The power consumption of the thruster is put at

10𝑊 which is also on a high side for CubeSats. Aside from the power requirement, the

major disadvantages of the cold gas thrusters are leakage at the valves and connectors,

clogging problems, heavy and high-pressure propellant storage which equally decreases

the propellant mass fraction [4, 67, 71], which makes the cold gas thruster unsuitable

for de-orbiting purposes. Also, the restriction of 1.2 𝑎𝑡𝑚 (0.12159𝑀𝑃𝑎) [1] pressure

regulation on board a CubeSat for launch has made a pressurised gas much less attractive

option for CubeSat applications, and their low specific impulse limits their use for orbital

transfer.

The use of MEMS technology to develop cold gas thrusters, and the replacement of

conventional valves with piezovalves have eliminated some of the associated problems

except the propellant storage under high pressure [4]. This problem was solved by

Rackemann 𝑒𝑡 𝑎𝑙 [72] that used solid gas generator cartridges, shown in Figure 1.19,

to store solid propellant. The propellant is only ignited when required to produce the

gaseous nitrogen to generate thrust. Power budget for the ignition process is put at

approximately 2.5𝑊 for 30 𝑠𝑒𝑐𝑜𝑛𝑑𝑠 [72], requiring energy that nanosatellites can not

afford.

19


Figure 1.19: Gas generator cartridges [72]

1.2.2.2 Warm Gas Thruster

In a quest to increase the efficiency of a cold gas thruster a heat exchanger is normally

placed before its nozzle and the new architecture is known as a warm gas thruster. The

increased temperature guarantees a higher specific impulse, with less propellant mass

producing the same thrust and thus higher efficiency [73]. Warm gas thrusters can

circumvent some of the limitations of cold gas thrusters by carrying propellant as a liquid,

and heating the liquid to a two phase state at elevated pressure before firing.

Figure 1.20: Schematic of a novel warm gas propulsion system [74]

French [75] designed a warm gas thruster for small satellites where he recorded about

50% improvement in thrust performance in excess of cold gas performance. The increase

in performance comes with a price of additional power, making the technology expensive

20


for nanosatellites. Shown in Figure 1.20 is a novel warm gas pressurization system that

was designed by Primex Aerospace Company (PAC). The system, which uses liquid

pressurant fuel, boasts of limited input power supply only for a short time for ignition

but operates on a pressure range of 810 𝑝𝑠𝑖𝑔 +5%/-7% [74], which is far beyond the

pressure regulation for nanosatellites.

Away from the warm gas systems are the hot gas systems whose source of heat is

self-generated as the propellant reacts with a suitable catalyst or with another propellant

to produce hot exhaust. The hot exhaust gas is a product of a chemical reaction that is

characterised by the combustion of propellants within the combustion chamber, and the

combustion products accelerated through a converging-diverging nozzle [71] to a high ve-

locity to create thrust. The propellants may be liquid, solid or both (hybrid). The liquid

propellant system (or Liquid Rocket Engines) is further classified into monopropellant or

bipropellant systems.

1.2.2.3 Monopropellant Systems

This is a single propellant system whose propellant is stable at ordinary atmospheric

conditions but decomposes exothermally into its constituents of hot gases when it passes

through a suitable catalyst. The heated high pressure gases are then expelled through

the nozzle to generate thrust with performances that exceed the cold and warm gas

thrusters, though with additional complexity and high temperature [76, 77]. The specific

impulse ranges of typical monopropellant systems are from 165 𝑠 - 244 𝑠 [78]. Hydrazine

and high concentrated (≥ 80 𝑏𝑦 𝑤𝑡) hydrogen peroxide, also known as High Test Peroxide

(HTP), are the heritage propellants with high performances that are commonly used for

monopropellant thrusters [76, 12, 79]. Hydrazine thruster is 20% higher in performance

than the HTP thruster, but the propellant is toxic, carcinogenic and flammable and

it requires special training for handling procedures and precautions [76, 4, 80]. This

leads to safety issue in propellant handling, which increases the total budget of a

nanosatellite mission. A hydrogen peroxide monopropellant thruster for nanosatellites

has been developed in Austrian Research Centres Seibersdorf by Scharlemann 𝑒𝑡 𝑎𝑙

[81, 82]. Figure 1.21 is one of several designs where they used hydrogen peroxide of

75%-87.5% concentration, and tested on different catalysts and mass flow rates. The

thruster, which operates on limited power requires no pre-heating of the catalyst bed

21


reducing its transition time to be about 10 𝑠𝑒𝑐. Recorded performance of the thruster

when operated at atmospheric condition indicates a thrust value ranging from 50 to

550𝑚𝑁 at a specific impulse of between 70 to 100 𝑠 [81].

Pressure andtemperature

gauges

Pressure andtemperature

gauges

Figure 1.21: Model achitechture of a miniature hydrogen peroxide monopropellantthruster, with a cross sectional view of the catalyst assembly [82]

HTP monopropellant is non-toxic and a cheaper alternative to hydrazine systems though

with lower performances. Also, a long storage of the propulsion grade hydrogen peroxide

as propellant will turn it into a dilute peroxide due to self-decomposition of the propellant

that also result in pressurisation due to oxygen evolution with a significant pressure

[83, 12, 79]. The decomposition rate which increases with temperature (at about 2.3

times per 10∘C rises) will drastically affect the performance of the propulsion system

against the targeted performance. Also, the high temperature generated through the

decomposition of the propellants when in contact with a catalyst bed creates thermal

control issues for nanosatellites.

1.2.2.4 Bipropellant Thrusters

Bipropellant systems use two separate tanks with different delivery systems to store and

inject both the fuel and the oxidiser, such as hydrogen and oxygen, into the combustion

chamber for chemical combustion. Just like the monopropellant system, the combustion

products are then directed through a converging-diverging nozzle to generate thrust.

They are mostly found on larger satellites for primary propulsion applications where

22


high impulse thrust is required, with monomethyl hydrazine/nitrogen tetroxide thruster

being the most common choice for in-space propulsion [84], though with high toxicity

levels [82]. These propellants raise safety concern and cost on their applications on

nanosatellites. Bipropellant systems typically generate higher levels of thrust than what

is normally required for nanosatellites. But in recent years, MIT has developed micro

bipropellant engine from a stack of silicon wafers using MEMS-based technology [85] as

shown in Figure 1.22. The thruster system measuring 18𝑚𝑚× 13.5𝑚𝑚× 3𝑚𝑚 is an

integration of combustion chamber, turbine pumps, inlet valve and the nozzle [ 4]. The

high-pressure thruster has demonstrated 1𝑁 of thrust with a thrust power of 750𝑊

and at a chamber pressure of 12 𝑎𝑡𝑚. Bipropellant systems are generally complex for a

nanosatellite mission.

UN

CO

RR

Figure 1.22: Micro-bipropellant thruster from MIT [85]

1.2.2.5 Solid Rocket Motor

Solid rocket motors use a solid propellant mixture called grain for a one shot combustion.

The propellant, which contains both the fuel and oxidiser is stored in the combustion

chamber, and the hot gas from the combustion is accelerated through a hollow cavity

within the grain once the grain is ignited to generate thrust [12, 71]. The major advantages

of solid rocket motors over the liquid rocket engines are their simplicity, storability and

the size of volume the propellant occupies by the same propellant mass. However, there

is yet no mechanism to stop the burning once ignited. In view of this setback, engineers

in Laboratory for Analysis and Architecture of Systems (LAAS CNRS) in France came

up with the concept of solid propellant thruster (SPT) in 1997 [86]. Though the principle

of operation is still on a one shot basis for a high rate of combustion, several arrays of

23


SPTs are fabricated in a micromachined silicon chip for multiple shots. A single chip of

24𝑚𝑚 × 24𝑚𝑚 dimension contains 10 × 10 single SPT, making a total of 100 SPTs.

Each SPT contains an igniter, a propellant reservoir, a nozzle and a seal wafer, as shown

in Figure 1.23. The propellant reservoir is packed with glycidyle azide polymer mixed

with ammonium perchlorate and doped with tiny particle of zirconium (GAP/AP/Zr) or

a more sensitive and energetic zirconium perchlorate potassium (ZPP) [87].

1,5mm

30

0µ

m3

60µ

m1

mm

Seal part

Reserviors part

Igniters part

Nozzle part

Figure 1.23: Schematic view of one solid propellant thruster [86]

Early performance of SPT has shown a total impulse of 1.5𝑚𝑁𝑠 and a thrust of about

5𝑚𝑁 from firing a cavity that has a throat diameter of 110𝜇𝑚, chamber size of 850𝜇𝑚

diameter and 1𝑚𝑚 length, over a duration of 500𝑚𝑠 [4, 38]. The major setbacks of SPT

are creation of debris particles around the throat section and within the satellites due

to incomplete combustion, poor repeatability of cavity ignitions, uncontrolled rupture

of neighbouring cavities due to heat loss [87]. A similar approach has been developed

from a joint research by the Aerospace Corporation, TRW and California Institute of

Technology [88], with a difference of the heating point of propellant in the cavity [38].

Known as digital micropropulsion, each layer of the thruster is a sandwich of three layers

containing micro-resistor, thrust chamber and a rupture diaphragm, and a complete

digital thruster could contain up to 10 6 thrusters. Their initial test, using a styphnate

propellant has produced a total impulse of 10−4𝑁𝑠 but at an input power of 100𝑊

[88].

24

1.3. CubeSat Requirements for Propulsion System

1.2.2.6 Hybrid Propulsion System

A hybrid propulsion system is combination of solid and liquid propellant engines with

hybrid traits. The fuel, in solid phase, is stored in the combustion chamber while the

oxidiser in either liquid or gaseous phase is stored away from the chamber. The inlet of the

oxidiser into the chamber necessitates a chemical reaction with the pyrolysed gaseous fuel

whose products are characterised with high temperature (in the range of 500∘C to 800∘C)

and pressure [89]. The products are accelerated through a converging-diverging nozzle to

create thrust. Figure 1.24 shows the schematic diagram of a convectional hybrid rocket

motor. The associated high temperature during its operation is far beyond a typical

payload temperature range of a nanosatellite (-40 to 85 ∘ [90]). However, it has certain

features that make it advantageous over either liquid or solid types of thrusters: low cost,

and simple, and the operation can be stopped and restarted, reliable with throttleable

thrust levels. We shall explore these advantages for CubeSat applications.

Figure 1.24: Schematic of a conventional hybrid rocket motor [91]

1.3 CubeSat Requirements for Propulsion System

The on-board power, mass, volume and size of a CubeSat preclude the use of most of the

micropropulsion systems whose input power is beyond 1𝑊 and/or takes more than half

its mass and volume for propulsion system. The current propulsion systems as listed in

Section 1.2 are analysed against these requirements as shown in Table 1.3 for performance

comparison and for CubeSat propulsion requirements. For example, the limited on-board

25


power system restricts the use of high fuel-efficient electric propulsion systems that is

power dependent while the energy dependent chemical propulsion requires a larger mass

and volume of the CubeSat for its propellant for higher ΔV requirements.

Table 1.3: Performance comparison of micropropulsions as requirements for CubeSatpropulsion system

Propulsion F 𝐼𝑠𝑝 𝐼𝑡𝑜𝑡* P 𝜌 𝑀𝑚 F/P* 𝑚𝑝** Burn

type (mN) (s) (Ns) (W) (𝑘𝑔/𝑚3) (g/mol) (mN/W) (kg) time* (s)

Electric

Electrothermal

Micro-Resistojet

Water 0.305 100 491 7.9 1000 18.02 0.039 0.500 447 hrs

Butane 0.01-0.1 95-100 269.3 17.5 578 58.12 0.57 0.289 7.48 hrs

Micro-Arcjet

Ammonia 1.2 147 491 6 0.73 17.03 0.2 0.00037 7.4 mins

Electrostatic

Ion engine-Xenon 0.059 5,480 161 13 5.894 131.3 0.005 0.003 31.7 days

𝜇 FEEP-Cesium 0.0001-0.15 3000-45000 35,664 6 1930 168.3 0.013 0.965 5.5 days

Hall thruster-Xenon 7.3 940 27.7 10 5.894 131.3 0.0007 0.003 1.05 hrs

Electromagnetic

𝜇 PPT-Teflon 0.06-0.275 266 2,873.3 15.6-36 2200 350000 0.000006 1.100 198.5 days

Cold gas

𝑁2 0.001-10 65 0.4 10 1.251 28.01 0.0005 0.00062 1.32 mins

Butane 10-25 70 0.9 1 2.48 58.12 0.0125 0.00124 68.2 secs

Chemical

Monopropellant

𝐻2𝑂2 50-550 65 459.6 10 1440 34.01 30 0.72 1.53 secs

Bipropellant

𝐻2 +𝑂2 2000 266 0.71 18 1.518 34.01 111.11 0.00027 0.35 secs

Solid motor

𝐶𝑜𝑚𝑝𝑜𝑠𝑖𝑡𝑒 𝑚𝑎𝑡𝑒𝑟𝑖𝑎𝑙1 0.1 0.1 150 0.0007

* calculated values, ** assuming the propellant mass occupies half a volume of 1𝑈 CubeSat, 1 Glycidyle azide polymer mixed with ammonium

perchlorate and doped with tiny particle of zirconium,F=Thrust, 𝐼𝑠𝑝=specific impulse, 𝐼𝑡𝑜𝑡=total impulse, P=electric power input, 𝜌=density,

𝑀𝑚=molecular mass, 𝑚𝑝=propellant mass

Though the electric propulsion systems have higher total impulse, the lower thrust values

is reflected in the high burn time for orbital manoeuvre even reaching months in some

cases, and a normalised input power to 1𝑊 will cause the manoeuvring time to reach

years. However, chemical propulsion systems operating at low power provides higher

thrust values and shorter orbital manoeuvre time in the range of seconds.

1.3.1 Propulsion Requirements for CubeSat Missions

Conventional satellite with propulsion systems on-board performs different orbital adjust-

ments to reposition itself once launched into space depending on the its mission. These

26


orbital adjustment include attitude control for the satellite to control its orientation

for precise nadir pointing of its payload and detumble the angular rate of the satellite;

orbital maintenance which helps to prolong the satellite mission life by counteracting

atmospheric drag especially at lower altitude; and deorbting the satellite after the mission

life into a parking orbit or grave yard to mitigate against space debris. Expanding the

capability of CubeSat will require the nanosatellite to perform such orbital manoeuvra-

bility as the conventional satellites, and therefore need be equipped with propulsion

system. According to Perez 𝑒𝑡 𝑎𝑙 [92], the propulsion requirements for nanosatellites for

several missions include the ability of the propulsion system to deliver ΔV between 1 to

100𝑚/𝑠. Other requirements are summarised in Table 1.4.

Table 1.4: Propulsion requirements for nanosatellites for several missions [92]

Parameter description Nanosatellite propulsion requirements

Thrust level range, 𝑚𝑁 1 - 1000

Micro-thruster wet mass, 𝑘𝑔 3

Micro-thruster power consumption, 𝑊 10

Micro-thruster volume, 𝑚3 0.0008 - 0.009743

Micro-thruster lifetime, 𝑦𝑒𝑎𝑟𝑠 2 - 5

Number of micro-thruster per nanosatellite 1 - 12

Minimum impulse bit, 𝑚𝑁𝑠 0.1 - 100

Judging from the wet mass of the micro-thrustre, the size of the nanosatellite put

additional constraints the these requirements. For instance, a 1.33 𝑘𝑔 CubeSat with total

on-board power of 1.6𝑊 and can not carry a propulsion system whose net mass is 3𝑘𝑔

and consumes an average power of 10𝑊 . A CubeSat will need a propulsion system that

operate within the on-board resources and yet deliver enough ΔV for different mission

scenarios. Depending on the requirements however, fuel efficient electric propulsion is

needed where larg ΔV and low thrust are required for long period orbital change. But

in some cases, an impulsive orbital manoeuvre is preferred, which is an attribute of

chemical propulsion system though with lower ΔV.

27


1.3.2 Chemical Propulsion Trade-off

Though the chemical propulsion systems have potential applications on nanosatellites in

term of low input power, there are other factors that are attributed of specific system as

seen in Table 1.5, which require attention to make the chemical system a viable option

for CubeSat applications. For example, a cold gas system though simple requires heavy

tank to withstand the pressurised gas and in most cases, the vapour pressure of the gas

exceeds the regulated pressure for a CubeSat propulsion system. This is the case of a

butane cold gas thuster whose vapour pressure is about 2 𝑏𝑎𝑟 at 21∘C. Other factors

affecting the choice of other chemical propulsion systems are shown in Table 1.5.

Table 1.5: Chemical propulsion trade-off for CubeSat applications

Monopropellant Bipropellant Solid rocket motor Hybrid rocket motor

1. Hybrazine system -Generally complex -One-way shot -Non-storable oxidiser

-Toxic substances for CubeSat (no mechanism to -Scaling issue

-Carcinogenic applications stop the burning onces -Thermal control problem

-Flammable ignited)

-Requires special training

for handling procedures and Merits of hybrid system

precautions over either liquid or solid

2. Hydrogen peroxide system types of thrusters:

-Self-degradable of propellant -Low cost and simple

over time -The operation can be restartable

-Performance reduction -Reliable with throttleable

-Thermal control problem thrust level

In view of the above hindrances of liquid and solid propulsion systems, one can leverage

on the benefits of hybrid system to design a low cost, simple,restartable and reliable

propulsion system for CubeSat applications.

1.3.2.1 Overview of Hybrid Rocket Motor

The development of hybrid rocket motors started in the 1930𝑠 with a focus mainly on

launch applications [93, 12]. The first research was conducted by S. P. Korolev and

M.K. Tikhonravov using gaseoline-collophonium mixture and liquid oxygen on a 500𝑁

thrust motor which was tested in 1933 to propel a rocket to an altitude of 1500𝑚

[94]. Among other applications of this technology in literature are sounding rockets,

28


target drones, tactical motors, and specific space applications like the transfer stage

and airborne launcher as reviewed by Frota and Ford [95]. The development and use

of hybrid rocket motors as microsatellite propulsion systems are being pursued around

the world. For example, SpaceDev in August 1999 was awarded a contract to develop a

micro-kick hybrid motor that is storable, re-startable, throttleable, modular and scalable

for microsatellites. The motor has 130𝑚𝑚 diameter and 305𝑚𝑚 length with a total

thrusting time of about 45 𝑠 [96, 91]. Also ONERA in France is developing a hydrogen

peroxide (H2O2)/polyethylene or Hydroxyl-Terminated Polybutadiene (HTPB) hybrid

propulsion system for 100𝑘𝑔 microsatellites and small tactical missiles [96]. At present,

the attention on the choice of propellants has been focussed on HTPB, Poly-Methyl

Methacrylate (PMMA) and Polyethylene (PE) as fuel and Liquid oxygen (LO), H2O2

and Nitrogen oxide (N2O) as the oxidisers [89, 97, 71]. Though these combinations

provide over 300 𝑠 vacuum specific impulse with high storage density, the associated heat

transfer during combustion processes is high and some of these oxidisers are not storable

for space applications [12, 95]. Other contending issues are low combustion efficiency,

higher sliver fraction and low regression rate of solid fuel. Over the years, there have been

research efforts to perfect the technology in terms of improving the regression or burning

rate [89, 98, 97] and making the technology suitable and applicable to microsatellites

[99] by:

∙ chemical methods which involves the preparation of the solid fuel with additives

like AP (Ammonium Perchlorate) and Al (Aluminium) powder to increase the

production of heat in the solid fuel thereby increasing the regression rate

∙ physical methods in which the geometry of the fuel is altered and the position and

orientation of the oxidiser inlets are adjusted. This includes embedding metal wire

into the solid fuel to increase the burning rate of the fuel and the use of swirl flow

of the oxidiser for the enhancement of regression rate [100]. For example, Haag [99]

in 2001 incorporated a vortex injection scheme (see Figure 1.25), which provides a

swirl flow of the oxidiser into the fuel grain of the Hybrid Rocket Motor (HRM) to

increase the performance over the conventional HRM and make the design suitable

for small satellite applications.

29


(a) (b)

Figure 1.25: (a) Vortex flow pancake hybrid model diagram [76] (b) Swirling of propellantin a vortex flow [101]

The current propellants selections for hybrid system precludes their viability for CubeSat

application despite their associated system advantages over liquid and solid counterparts

as shown in Table 1.5. However alternative propellants whose enthalpy of reaction and

the reaction temperature are within structural limits of the CubeSat are needed while

employing the benefits of the system. Therefore a thermochemical analysis of propellants

combinations is performed using NASA Chemical Equilibrium with Applications (CEA)

Program, a computer program for theoretical rocket performance from chemical combi-

nation of propellants through equilibrium compositions of their constituent mixtures and

reaction products [102, 103], to determine their transport and thermodynamic properties

of the individual propellant and their species, and quantify them for CubeSat propulsion.

The propellants analysis is shown in Table 1.6.

30


Table 1.6: Thermochemical analysis of propellant combinations

S/N Propellants Ratio Pc Tc 𝜌 Specific enthalpy 𝐼𝑠𝑝 Δ𝑉 *

(𝑏𝑎𝑟) (𝐾) (𝑘𝑔/𝑚3) (𝑘𝐽/𝑘𝑔) (𝑠) (𝑚/𝑠)

1F (Polyethylene(s))

5.55 44.7 3091.3855 -2704.0

325.3 412.2O (𝐻2𝑂2 (l)) 1580 -160815.7

2F (Polyethylene(s))

5.55 44.4 4609.8855 -2704.0

402.4 509.9O (𝑁2𝑂(l)) 808.6 431694.1

3F (Al(s))

5.55 44.7 3771.62800 189703.8

258.3 327.3O (𝑂2(l)) 1140 -12979.0

4F (Al(s))

5.55 44.4 1601.52800 189703.8

224.4 284.3O (𝐻2𝑂(l)) 1000 -260446.8

* calculated ΔV values assuming 88% dry mass ratio for 1𝑈 CubeSat, 𝑃𝑐 is the chamber pressure,

𝑇𝑐 is the chamber temperature, 𝜌 is the propellant density and 𝐼𝑠𝑝 is the specific impulse of the

combination

The propellants combinations are treated to the same ratio and nozzle expansion ratio

of 50. The result reveals that the hydrocarbon combustion reaction with hydrogen

peroxide and nitrogen tetra-oxide have higher specific impulse and ΔV for the same dry

mass ratio than the aluminium oxidation reaction. However, while hydrogen peroxide

self degrades over time and thus losses performance, nitrogen tetra-oxide is a harmful

and toxic gas that requires special training for handling procedures and precautions.

These combinations are well suitable for operation that does not require long storage

of propellant, especially during the launch operation. Another point to consider is the

chamber temperature. Thermal control is associated with the conventional hybrid rocket

motor where both hydrogen peroxide and nitrogen tetra-oxide are currently used as

oxidisers. But a prerequisite condition of a CubeSat propulsion system is to operate

within the structural temperature of the satellite thereby reducing the cost in employing

mechanism to control thermal runaway in the system. Therefore aluminium oxidation

reaction is considered as propellants combination choice. In this regard, both Al/LO and

Al/𝐻2𝑂 reactions are considered. While liquid oxygen is cryogenic in nature and it has

a boiling point of -183∘C, and requires special training and equipment for handling and

storage[104], liquid water is easy to store and has a boiling point of 100∘C. This reaction

is also known as water splitting reaction, which is normally used to generate hydrogen

gas. Therefore only liquid water oxidation reaction with aluminium is considered as

alternative propellant for CubeSat propulsion system in this research. This propellant

31

1.4. Motivation and Objectives

combination has theoretical performance of specific impulse of 224.4 𝑠 and is capable

of delivering a ΔV of about 284𝑚/𝑠 for a 1𝑈 CubeSat of 1.33 𝑘𝑔 assuming dry mass

fraction of 88%.

1.4 Motivation and Objectives

Until quite recently CubeSats did not have propulsion systems onboard and their orbits

were dictated by the injection orbit supplied by the launch vehicle (LV). That was

primarily due to the fact that CubeSats applications were for experimental purposes

within the university communities and were done on extremely low budgets. The involve-

ment of government establishments and private organisations in CubeSat applications

has brought the need to extend the capabilities and altitude of the CubeSat beyond

low earth orbit with the development of low-cost propulsion systems that is within the

power, size and mass of the nanosatellites thereby making the nano-group satellites

serious contenders for significant science missions. Missions beyond earth observation

to technology demonstration around other celestial bodies like the asteroid belt and

the outer solar system require significant orbital change and ΔV capability, as well as

ΔV requirements to de-orbit the satellite after the mission assignment. Available micro-

propulsion systems are considered and reviewed in the next section for a 1𝑈 CubeSat

against its limited on-board resources. Though electric propulsion systems are effective

in fuel efficient but these systems are complex and they put a power-demand burden on

CubeSat for moderate to high thrust applications. Chemical propulsion systems with

less fuel efficient could provide impulsive orbital change for CubeSat missions, but the

current propellants are toxic, carcinogenic, flammable, and they require special trainings

for handling procedures and precautions, and this will result in high mission cost. And

also, simple and cheap cold gas thruster loses its performance when used for end of

life de-orbiting purpose. Among the propellants combinations for a hybrid system for

CubeSat propulsion system, only Al/𝐻2𝑂 is chosen considering the ease of handling

and storing the propellant. Therefore a simple alternative system is investigated in this

study whose propellants are readily available and cheap for CubeSat budget, and whose

performance remains the same through the life span of the nanosatellite.

A propose propulsion system in this research is a high density propellant (alumini-

32

1.5. Novelty and Research Achievements

um/sodium hydroxide/water) hybrid propulsion system for CubeSat applications. It

involves the injection of liquid oxidiser (mixture of sodium hydroxide and water) into a

high density solid fuel (aluminium wool) which will produce water vapour and hydrogen

gas for thrust generation. Unlike a conventional hybrid rocket motor, the chemical

reaction is not a high temperature combustion process. On the contrary, this reaction is

self-sustaining and progresses at moderate temperature of about 150∘C, and therefore

scalable without thermal and combustion instability issues [41]. The materials are

storable, low cost and readily available, with aluminium having a high energy density

per unit volume. The propulsion system only requires minimal power, about 500𝑚𝑊 ,

from the CubeSat electrical bus system for valves operation. One of the drawbacks of

the system is that the firing time is about 99 𝑠 long, making this system inappropriate

for attitude control purposes.

1.5 Novelty and Research Achievements

The areas of novel contributions arising from the course of this work to the field of

micropropulsion systems for CubeSats include:

∙ The novel use of the chemical combination of aluminium wool/water/sodium

hydroxide as propellant for CubeSat propulsion.

∙ Developing a new CubeSat propulsion architecture that could potentially deliver

high ΔV capability for future CubeSat missions.

33

Chapter 2

Theory

2.1 Overview

This section reviews the reaction between aluminium, water and sodium hydroxide for the

generation of hydrogen gas, and shows how this reaction can be adopted for nanosatellites

propulsion system. It also discusses thermodynamics and dynamics of gas with suitable

assumptions for propulsion systems. Finally, it contains a o-dimensional propulsion

model and the ideal performance resulting from the analysis.

Water splitting reaction with reactive metals has been a subject of research for the

generation of hydrogen gas in automobile industries [105, 106, 107]. The reactive metals

have high energy densities such that a very small amount is suited for volume constraint

nanosatellites propulsion system. This exothermic reaction is adopted in this research

and among the metals (zinc, magnesium, calcium, aluminium, e.t.c) from the available

data in literature in term of cost, safety, availability, aluminium has the best bargain [107].

Aluminium is a very reactive metal that when placed into the atmosphere reacts quickly

to form a film of aluminium oxide on the surface of the metal. Though the aluminium

oxide film is microscopically thin, in the range of 50 to 100Angstroms [108], and gener-

ally very stable under the pH scale of 4 - 9 [109], it protects the surface of the metal

from further reaction with the surroundings thereby preventing the potential use of the

metal for energy application purposes except when treated. But when the thin layer

or barrier is broken down, the energy stored in this abundant metal becomes useful.

34

2.1. Overview

There are two basic oxidation methods of aluminium in the literature: aluminium -

oxygen reaction and aluminium - water reaction. The first is a combustion reaction

involving aluminium powder in the presence of an oxidiser with uncontrollable release

of heat energy. Aluminium powder is used as a solid propellant fuel in aerospace ap-

plications [110, 111] and in pyrotechnics [112]. The second oxidation method is a low

temperature, non - combusting and controllable reaction used in the generation of energy

and hydrogen gas. [113, 114, 115, 116, 117, 118]. Shkolnikov 𝑒𝑡 𝑎𝑙 [119] summarised

the different methods that is used to increase the efficiency of aluminium based energy

generation technology, which include alloying with certain elements, chemical activator

addition, mechanical processing or mechanochemical treatment and heating. Equation

2.1 shows the reactions and the energy applications of the aluminium - water oxidation

reaction. The enthalpy of the reaction calculated from the enthalpy of formations in

Table 2.1 (data obtained from 𝐶𝑅𝐶 𝐻𝑎𝑛𝑑𝑏𝑜𝑜𝑘 𝑜𝑓 𝐶ℎ𝑒𝑚𝑖𝑠𝑡𝑟𝑦 𝑎𝑛𝑑 𝑃ℎ𝑦𝑠𝑖𝑐𝑠 [120]) is

419 𝑘𝐽/(𝑚𝑜𝑙𝑒 𝑜𝑓 𝑎𝑙𝑢𝑚𝑖𝑛𝑖𝑢𝑚), with a release of about 3 𝑔 or 1.5𝑚𝑜𝑙𝑒𝑠 of hydrogen gas.

Figure 2.1 shows the relationship between the reacted moles of aluminium and the moles

of released hydrogen gas and the enthalpy of the reaction

1 2 3 4 5 6 7

Mole of aluminium (mol)

0

2

4

6

8

10

12

Mo

le o

f h

ydro

ge

n g

as

(mo

l)

1 2 3 4 5 6 7

Mole of aluminium (mol)

0

500

1000

1500

2000

2500

3000

En

tha

lpy

of

rea

ctio

n (

kJ)

Figure 2.1: Moles of hydrogen gas produced and the enthalpy of reaction against themoles of reacted aluminium respectively

That means when a fraction of a mole of aluminium, say 0.05𝑚𝑜𝑙𝑒𝑠, reacts with water in

the presence of favourable promoter, about 20.95 𝑘𝐽 of heat will be released and about

0.075𝑚𝑜𝑙𝑒𝑠 of hydrogen gas will be produced. Again, when 1 𝑘𝑔 of aluminium reacts

with water in the presence of favourable chemical promoter or activator, about 15 -

16𝑀𝐽 of heat will be released, 0.111 𝑘𝑔 of hydrogen gas will be produced. All these are

based on complete reaction of the reactants to form the products. The applications of

35

2.2. Aluminium, Sodium Hydroxide, Water Oxidation Reaction

the reaction include: on - demand production of hydrogen gas for portable power sources

and stationary power plants for electricity generation; source of heat energy for heating

purposes; the by - product, aluminium hydroxide, is used as aluminium salt for water

treatment and as hydroxide in pharmaceuticals, and it can be recycle to recover the

aluminium [116].

𝐴𝑙(𝑠) + 3𝐻2𝑂(𝑙) −→ 𝐴𝑙(𝑂𝐻)3(𝑎𝑞) + 1.5𝐻2(𝑔) (2.1)

Table 2.1: Thermodynamic properties of the propellants

Substances Molar mass Enthalpy of formation Specific heat (𝑐𝑝)

kg/mol kJ/mol kJ/mol.K

𝐴𝑙(𝑠) 0.0269815 0 0.0242

𝐻2(𝑔) 0.0020159 0 0.028868

𝐻2𝑂(𝑙) 0.01801528 -285.8 -

𝐻2𝑂(𝑔) 0.01801528 -241.8 0.036031

𝐴𝑙(𝑂𝐻)3(𝑎𝑞) 0.0780036 -1276 -0.128706

2.2 Aluminium, Sodium Hydroxide, Water Oxidation Re-

action

Aluminium finds its use in space applications from the earliest days of space technology.

Spuknik 1, the first artificial satellite, was made out of aluminium [121]. The light weight

and low density (2700 𝑘𝑔/𝑚3) characteristics of the metal make it a useful component

in the body structure of spacecraft, providing a substantial weight reduction compared

to many other metals. Apart from its light weight, aluminium has a specific energy of

about 16𝑀𝐽/𝑘𝑔 when oxidised with water as shown in Equation 2.4 and it has a high

energy density of 83.8𝑀𝐽/𝐿 when oxidised with oxygen [122, 123]. This energy per

unit volume is highest than any other fuel that has been utilised for propulsion to date

(see Figure 2.2), making it a suitable choice in space propulsion applications as solid

fuel additive in both solid and hybrid rocket motors to enhance the regression rate as

well as the specific impulse of the fuels [98, 71]. Our reaction is based on the aluminium

oxidation by water.

36


Figure 2.2: Selected energy density of some fuels [122]

This high energy content can be accessed by reacting aluminium with water to form

its hydroxide or oxide, and hydrogen gas depending on the reaction temperature [106].

However, direct reaction of pure aluminium with water is inhibited by the formation of

a dense layer of Al2O3 which passivates the aluminium and stops any further reaction

[116, 119, 117, 123] thereby reducing its potential usage for energy and hydrogen genera-

tion.

Current efforts to increase the efficiency of the aluminium-water reaction include alloying

aluminium with a low melting point metals like gallium, indium, and tin [123], using

nanoparticle sizes of aluminium [124], addition of selected metallic salts [125], metal

oxide addition [126] and metal hydroxide addition [127, 115, 116]. In all these techniques,

oxidation of aluminium in alkaline solution to produce hydrogen occurs at moderate

temperature (about 150∘) and pressure [119], and among the alkaline additives, sodium

hydroxide is the best hydroxide for the reaction in view of reaction rate and conversion

37


degree [115, 116, 119, 117] if the associated corrosion of the system apparatus is miti-

gated. Hong-Bin Dai 𝑒𝑡 𝑎𝑙 [113] treated the sodium hydroxide corrosion with a small

amount of sodium stannate in aluminium-water reaction to generate hydrogen without

compromising the reaction rate in order to mitigate against corrosion.

The potential use of powdered aluminium/steam propellants and powdered alumini-

um/oxygen propellants for spacecraft propulsion were suggested by Ingenito [128] and

Ismail [110]. In their submissions, the reaction would involve igniting the combination to

a high temperature (about 2500K) to crack the aluminium protective layer for continued

reaction. The high temperature and ignition involved in the process necessitate tempera-

ture control measure and ignition systems which are complex for nanosatellite platforms.

In this study we took the approach of using a high energy density solid fuel-aluminium

wool (aluminium wool is preferred because it is in between highly inflammable nano-

particle aluminium when exposed to air and aluminium shavings that has less surface

area for reaction. Aluminium foil is also not chosen because we needed aluminium wool,

see 𝐴𝑝𝑝𝑒𝑛𝑑𝑖𝑥 𝐵 for the data sheet, to serve as filter for the reaction products before the

exhaust) and the combination of sodium hydroxide and water as oxidiser for a volume

constrained CubeSat as a hybrid propulsion system. This is a proof of concept design.

The materials are readily available, low cost and the reaction of the propellants occurs at

moderate temperatures, about about 150∘C. Though the operating temperature is less

than most chemical propulsion systems, the trade-off is in thermal control issue on other

subsystems on-board the nanosatellites, and the chemical reaction is non-combustible.

The sodium hydroxide acts to avoid the passivation of the aluminium and allows it to

react with water. The chemical reactions of aluminium, sodium hydroxide and water

occur in stages and are shown in equations (2.2), (2.3) and (2.4) [129, 115].

2𝐴𝑙(𝑠) + 6𝐻2𝑂(𝑙) + 2𝑁𝑎𝑂𝐻(𝑠) −→ 2𝑁𝑎𝐴𝑙(𝑂𝐻)4(𝑎𝑞) + 3𝐻2(𝑔) (2.2)

𝑁𝑎𝐴𝑙(𝑂𝐻)4(𝑎𝑞) −→ 𝑁𝑎𝑂𝐻(𝑠) +𝐴𝑙(𝑂𝐻)3(𝑎𝑞) (2.3)

2𝐴𝑙(𝑠) + 6𝐻2𝑂(𝑙) −→ 2𝐴𝑙(𝑂𝐻)3(𝑎𝑞) + 3𝐻2(𝑔) Δ𝐻∘𝑟𝑥𝑛 = −838𝑘𝐽 (2.4)

The reaction between water and sodium hydroxide occurs as a dissolution of sodium

hydroxide into aqueous sodium ion and hydroxide ion, which is an exothermic process

that releases heat energy [130]. Only water is consumed in equation (2.2) because of

38


the regeneration of sodium hydroxide from the decomposition of sodium aluminate in

equation (2.3), which acts as a catalyst [113, 115, 118]. Hence equation (2.4) is the

overall reaction. From thermochemical analysis of the Equation 2.4, the enthalpy of the

chemical reaction and the associated temperatures can be expressed using Equation 2.5

[131].

ℎ𝑐ℎ𝑒𝑚 = Δℎ∘𝑟𝑥𝑛 +

∫ 𝑇

𝑇𝑟𝑒𝑓

𝑐𝑝 (𝑇 ) 𝑑𝑇. (2.5)

The specific enthalpy of formation, Δℎ∘𝑟𝑥𝑛, is easily calculated using Hess’s law [132],

which states that the overall heat of reaction of a thermodynamic chemical reaction is

equal to the sum of the heat of formation of the reaction products minus the sum of the

heat of formation of the reactants. That is,

Δℎ∘𝑟𝑥𝑛 = ΣΔℎ∘𝑓(𝑝𝑟𝑡𝑠) − ΣΔℎ∘𝑓(𝑟𝑐𝑡𝑡𝑠) (2.6)

where ΣΔℎ∘𝑓(𝑟𝑐𝑡𝑠) is the specific standard heat of formation of the reactants and Δℎ∘𝑓(𝑝𝑟𝑡𝑠)

is the specific standard heat of formation of the products. Table 2.1 shows the Δℎ∘𝑓 values

of the propellants.The temperature limits, 𝑇𝑟𝑒𝑓 and 𝑇 , of the second term in Equation

2.5 are the reference temperature (it is normally 23∘𝐶) and the reaction temperature

respectively. The temperature-dependent heat capacity, 𝑐𝑝, is given in [133, 134] as;

𝑐𝑝(𝑇 ) = 𝑎+ 𝑏𝑇 + 𝑐𝑇 2. (2.7)

Therefore, ∫ 𝑇

𝑇𝑟𝑒𝑓

𝑐𝑝 (𝑇 ) 𝑑𝑇 = 𝑎𝑇 +𝑏𝑇 2

2+

𝑐𝑇 3

3

𝑇

𝑇𝑟𝑒𝑓

, (2.8)

where the constants 𝑎, 𝑏 and 𝑐 are obtained from [135, 136] and shown in Table 2.2.

These values are within the reaction chamber operating temperature.

39

2.3. Thermodynamics and Gas Dynamics

Table 2.2: Heat capacity coefficients

Substances Heat capacity coefficients from 300K to 1000K

𝑎𝐽/𝑚𝑜𝑙.𝐾

𝑏10−2𝐽/𝑚𝑜𝑙.𝐾2

𝑐10−5𝐽/𝑚𝑜𝑙.𝐾3

𝐴𝑙(𝑠)* 28.08920 -0.5414849 0.8560423

𝐻2(𝑔) 33.066178 -1.1363417 1.1432816

𝐻2𝑂(𝑙)** -203.6060 152.3290 -319.6413

𝐻2𝑂(𝑔)*** - - -

𝐴𝑙(𝑂𝐻)3(𝑎𝑞) 2.987191 10.7087 -8.95576* the range of temperature is from 298𝐾 to 933𝐾

** the range of temperature is from 298𝐾 to 500𝐾

*** no values at the operating temperature

Now, using Equations 2.4, 2.5, 2.6 and 2.8, and Tables 2.1 and 2.2, we can calculate

change in enthalpy of the reaction at 150∘𝐶, which is 418.6 𝑘𝐽/𝑚𝑜𝑙. This represents the

enthalpy of formation at 23∘𝐶 and the change in enthalpy at 150∘𝐶.

Roach 𝑒𝑡 𝑎𝑙 stated in their work that aluminium-water reaction produces a limited

amount of hydrogen gas because not all the aluminium takes part in the hydrogen-

formation reaction [137], which means the liberated energy in the reaction is less than

calculated above. This will be verified in the course of this research. However, if this

reaction is confined in a controlled volume with a small orifice, there will be a change of

water phase from liquid to gas due to low pressure environment and elevated temperature

of the reaction, and both the water vapour and hydrogen gas when directed through a

converging-diverging nozzle will produce thrust to propel the nanosatellites by expelling

the heated gas through a nozzle. The energy released in this reaction is about 838 𝐾𝐽

per 2𝑚𝑜𝑙𝑒𝑠 of aluminium.

2.3 Thermodynamics and Gas Dynamics

The first law of thermodynamics is a statement of energy conservation [ 138], and for a

control volume with one inlet and one exit, it is written as [139]:

𝑑𝐸𝑐𝑣

𝑑𝑡= 𝑐𝑣 − 𝑐𝑣 + 𝑖

(ℎ𝑖 +

𝑣2𝑖2

+ 𝑔𝑧𝑖

)− 𝑒

(ℎ𝑒 +

𝑣2𝑒2

+ 𝑔𝑧𝑒

)(2.9)

40


where 𝐸𝑐𝑣 is the energy of the control volume at time 𝑡 and 𝑐𝑣 and 𝑐𝑣 represent the

net rate of energy transfer by heat and work across the boundary of the control volume.

𝑖 and 𝑒 are inlet and outlet mass flow rate, ℎ𝑖 and ℎ𝑒 are the total specific enthalpies

of inlet and outlet flow of the control volume, 𝑣𝑖 and 𝑣𝑒 are the inlet and exit velocities of

the flow, 𝑧𝑖 and 𝑧𝑒 are the vertical measurements of the inlet and outlet from the ground

level and 𝑔 is acceleration due to gravity at the surface of the earth. The bracket terms

represent the enthalpy, kinetic and potential energy that accompany the rate of mass

flow in and out of the control volume as a function of time. Consider a control volume

attached to a nozzle for the thermodynamic analysis of an expanded gas as shown in

Figure 2.3.

AtAe

peAepa Control volume

ThroatNozzle

Figure 2.3: Control volume with an attached nozzle: 𝐴𝑒 is the exit area of the nozzle, 𝐴𝑡

is the throat area, 𝑝𝑎 is the ambient pressure and 𝑝𝑒 is the exit pressure

Assuming the products of the reaction in the reaction chamber experience no significant

change in potential energy, no shaft work or shear work done, and under adiabatic and

steady flow conditions the process is given by [138]

ℎ1 +1

2𝑣21 = ℎ2 +

1

2𝑣22 (2.10)

That is, the sum of the specific enthalpy and the specific kinetic energy remain constant

for a given flow [131]. In this equation ℎ is the specific enthalpy (𝐽/𝑘𝑔) and 𝑣 is the flow

velocity (𝑚/𝑠).

If the gas in the control volume is stationary, then its enthalpy is converted into kinetic

energy, and equation (2.10) becomes

ℎ𝑜 = ℎ+1

2𝑣2, (2.11)

where the enthalpy, ℎ, is obtained by combining temperature-dependent internal energy,

41


𝑢 and flow work, 𝑝𝑣, [131]. That is,

ℎ = 𝑢+ 𝑝𝑣. (2.12)

Differentiating equation (2.12) with respect to temperature and treating the gas as

perfect gas results in

𝑐𝑝 = 𝑐𝑣 +𝑅, (2.13)

where 𝑐𝑝 and 𝑐𝑣 are the specific heat capacity at constant pressure and constant volume

respectively and 𝑅 is the specific gas constant (𝐽/𝑘𝑔𝐾) [131]. The ratio of the heat

capacities (also known as the specific heat ratio) is given by

𝛾 =𝑐𝑝𝑐𝑣

(2.14)

Then, for an adiabatic expansion, equation (2.11) becomes

𝑐𝑝𝑇𝑜 = 𝑐𝑝𝑇 +𝑣2

2. (2.15)

That is

𝑐𝑝 (𝑇𝑜 − 𝑇 ) =𝑣2

2. (2.16)

Simply put, the change in enthalpy is the same as the gain in the kinetic energy of the

exhaust. 𝑇𝑜 is the stagnation temperature, which is equivalent to the temperature of the

gas when brought to rest adiabatically. The term 𝑣2/2𝑐𝑝 is the dynamic temperature

and 𝑇 is the temperature at the point of interest along the stagnation streamline. The

sonic or the acoustic velocity of the gas from the reaction chamber through the throat is

[131]:

𝑎𝑜 =

√𝛾𝑅𝑇𝑐

𝑀𝑚

=√𝛾𝑅𝑜𝑇𝑐

(2.17)

where 𝑅 is the universal gas constant (8314.5 𝐽/𝑘𝑚𝑜𝑙.𝐾), 𝑅𝑜 is the specific gas constant,

𝑇𝑐 is the chamber temperature (𝐾) and 𝑀𝑚 is the gas molecular mass (𝑘𝑔/𝑘𝑚𝑜𝑙). The

gas flow velocity, as observed by a stationary observer, is related to the acoustic velocity,

𝑎𝑜, through Mach number, 𝑀 , by [131]

𝑀 =𝑣

𝑎𝑜(2.18)

42


At the throat of the nozzle where the flow transitions from sub-sonic to super-sonic,

the Mach number is unity as the flow velocity equals the acoustic velocity. The Mach

number is greater than one towards the nozzle exit which results in supersonic flow.

Now, substituting equations (2.17) and (2.18) into equation (2.15) and replacing 𝑅𝑐𝑝

with 𝛾−1𝛾 results in

𝑇𝑜

𝑇= 1 +

𝛾 − 1

2𝑀2. (2.19)

This expression can be related to the stagnation pressure, 𝑝𝑜, and static pressure in the

nozzle assuming isentropic expansion to give:

𝑝𝑜𝑝

=

(1 +

𝛾 − 1

2𝑀2

) 𝛾𝛾−1

. (2.20)

Equation (2.20) can be re-written as

𝑝𝑒𝑝𝑜

=

(1 +

𝛾 − 1

2𝑀2

𝑒

) 𝛾−1𝛾

. (2.21)

The gas exit velocity (actual velocity), 𝑣𝑒, from the nozzle is obtained by solving equation

(2.15) and replacing 𝑐𝑝 by 𝛾𝑅/(𝛾 − 1). The result is:

𝑉𝑒 =

⎯ 2𝛾𝑅𝑇𝑜

(𝛾 − 1)

(1−

(𝑝𝑒𝑝𝑜

) 𝛾−1𝛾

)

=√𝛾𝑅𝑇𝑜

⎯ 2

𝛾 − 1

(1−

(𝑝𝑒𝑝𝑜

) 𝛾−1𝛾

) (2.22)

The actual exit velocity depends on the molecular and chemical attributes of the pro-

pellant, and on the expansion ratio of the thruster [140]. The specific impulse of the

gas, 𝐼𝑠𝑝, which also defines the efficiency of the propellant utilization in the chamber is

obtained by dividing the expression for the exit velocity by acceleration due to gravity

at the surface of the earth, which is 9.81𝑚/𝑠2:

𝐼𝑠𝑝 =

√𝛾𝑅𝑇𝑜

𝑔

⎯ 2

𝛾 − 1

(1−

(𝑝𝑒𝑝𝑜

) 𝛾−1𝛾

)(2.23)

43


The area ratio, 𝑒, of the nozzle in Figure 2.3 is given by the expression:

𝑒 =𝐴𝑒

𝐴𝑡=

1

𝑀𝑒

((2

𝛾 + 1

(1 +

𝛾 − 1

2𝑀2

𝑒

)) 𝛾+1𝛾−1

) 12

, (2.24)

where 𝑀𝑒 in the Mach number at the nozzle exit [131]. The mass flow rate of the gas

through the nozzle is:

=𝐴𝑡𝑝𝑜𝑎𝑜

𝛾

(2

𝛾 + 1

) 𝛾+12(𝛾−1)

, (2.25)

and𝑎𝑜

𝛾(

2𝛾+1

) 𝛾+12(𝛾−1)

= 𝑐* (2.26)

where 𝑐* is the characteristic velocity which helps to analyse the performance of the

propellants and the chamber performance away from the nozzle. Hence;

=𝐴𝑡𝑝𝑜𝑐*

. (2.27)

The throat area of the nozzle, 𝐴𝑡, when choked is obtained from equations (2.25) and

(2.17), and is given as;

𝐴* =

𝑝𝑜√𝛾.

√𝑅𝑇𝑜

𝑀

(2

𝛾 + 1

)− 𝛾+12(𝛾−1)

(2.28)

A steady thrust force, 𝐹 , generated by the exhausted gas at the exit of the nozzle from

Figure 2.3 is

𝐹 = 𝑉𝑒 + (𝑝𝑒 − 𝑝𝑎)𝐴𝑒, (2.29)

which is a sum of momentum thrust and pressure thrust. The velocity of the gas as it

expands through the nozzle exit, known as the effective exhaust velocity is:

𝑐 = 𝑉𝑒 +𝐴𝑒

(𝑝𝑒 − 𝑝𝑎) . (2.30)

Both the effective and actual velocities are equal when the ambient and exit pressures

are equal. The coefficient of the thrust 𝑐𝐹 that relates to the performance of the nozzle

is given by:

𝑐𝐹 =𝐹

𝐴𝑡𝑝𝑜

=𝑐

𝐴𝑡𝑝𝑜(𝑝𝑒 = 𝑝𝑎) .

(2.31)

44

2.4. Theoretical Performance Analysis of the Propulsion System

Combining equations (2.27) and (2.31) will also give the effective exhaust velocity

as:

𝑐 = 𝑐*𝑐𝐹 . (2.32)

The total impulse is related to the propellant mass, propellant exit velocity and specific

impulse by (2.33):

𝐼𝑡𝑜𝑡 =

∫ 𝑡2

𝑡1

𝐹𝑑𝑡 = 𝑚𝑝𝑣𝑒

𝑣𝑒 =𝐼𝑡𝑜𝑡𝑚𝑝

and 𝐼𝑠𝑝 =𝑣𝑒𝑔𝑜

(2.33)

where 𝐼𝑡𝑜𝑡 is the total impulse, and 𝑡1 and 𝑡2 are the time intervals for the thrusting

period.

2.4 Theoretical Performance Analysis of the Propulsion

System

This section analyses the design target performance and the efficiency of the propulsion

system. Some assumptions are made to obtain the design target parameters of the

propulsion system which were used in the model stage. These include:

∙ the chamber conditions are pressure, 𝑝𝑐 = 4 𝑏𝑎𝑟 (400𝑘𝑃𝑎) and temperature, 𝑇𝑐 =

150∘C (423.15𝐾). The chamber pressure is kept at this value (which is less than

the saturated vapour pressure of water at 150∘C) such that we can have substantial

component of the oxidiser in the vapour phase at the operating temperature

∙ water vapour and hydrogen gas are released from the reaction in 0.95:0.05 ratio.

This ratio is based on results of the analysis of series of experiments that were

performed before the design (detailed of the analysis is shown in 𝐶ℎ𝑎𝑝𝑡𝑒𝑟 4). This

ratio was used for the partial molar mass of the exhaust gases and the specific

heat ratio, which were calculated to be 17.215 𝑘𝑔/𝑘𝑚𝑜𝑙 (molar mass of water

and hydrogen gas are 18.015 𝑘𝑔/𝑘𝑚𝑜𝑙 and 2.015 𝑘𝑔/𝑘𝑚𝑜𝑙 respectively) and 1.324

(interpolated specific heat ratio for water vapour at 150∘C is 1.317 and the specific

heat ratio for hydrogen gas is 1.4) respectively.

45


∙ the gases are treated as ideal gases

∙ the gases expand isentropically through a choked nozzle (Mach number is 1 at the

nozzle throat) in a steady flow

∙ for an impulsive chemical propulsion system, a thrust-to-spacecraft mass ratio of

0.22 is assumed and in compliance with similar designs in literature [131, 41, 12, 4,

56] , which is also ideal to de-orbit the nanosatellites. This ratio will give a thrust

of 0.29𝑁 for a 1.33 𝑘𝑔 CubeSat.

We therefore used the equations in Section 2.3 for the analysis. The acoustic velocity

of the gases from the chamber through the throat is obtained from Equation (2.17) as

520.18𝑚/𝑠. This value was used in Equation (2.26) for the characteristic velocity which

is 673.166𝑚/𝑠. The exhaust gases specific impulse, exit pressure, exhaust Mach number

and the cross sectional area of the nozzle exit are obtained from Equations (2.21), (2.23),

(2.24) and the specific impulse is plotted as a function of expansion ratio in Figure 2.4(a).

In this plot, there is sharp increase in the specific impulse performance at an expansion

ratio of 20 which begins to stabilise before 120𝑠 at an expansion ratio of 40. Hence the

expansion ratio of 40 was chosen. Figure 2.4(b) is a plot of the propellant specific impulse

as a function of mass flow rate, from Equation (2.29) using the desired thrust value and

assuming that the exit pressure equals the vacuum pressure. The plot shows a downward

trend in specific impulse as the mass flow rate of the propellant from the nozzle increases.

This shows that the utilisation efficiency of the propellant is inversely proportional to

its mass flow rate. However, propellant exhaust mass flow rate of 0.00025 𝑘𝑔/𝑠 from

the nozzle that correspond to the specific impulse of 118 𝑠 was used for the propulsion

design. This value was used to calculate the throat area of a choked nozzle from Equation

(2.28).

46


0 20 40 60 80 100 120 140 160

Expansion Ratio, 0

0

20

40

60

80

100

120

Spe

cific

Impu

lse,

Isp

(s)

X: 39.91Y: 118.4

(a) Specific impulse against expansion ratio(𝐴𝑒/𝐴𝑡)

2 2.5 3 3.5 4 4.5 5 5.5 6 6.5

Mass flow rate, kg/s # 10-4

40

50

60

70

80

90

100

110

120

130

Sp

eci

fic Im

pu

lse

, I s

p(s

) X: 0.0002528Y: 118

(b) Specific impulse against propellant mass flowrate

Figure 2.4: Theoretical specific impulse performance against the expansion ratio of thenozzle and the propellant mass flow rate

The design target initial ΔV performance of the propulsion system was obtained using the

Tsiolkowski equation, shown in Equation 2.34. Results of some pre-design experiments

show that the propellant mass fraction is 12%. The relationship is plotted in Figure

2.5 where a total ΔV of about 150𝑚/𝑠 at a specific impulse of 118𝑠 is obtained from a

propellant mass fraction of 12%.

Δ𝑉 = −𝑉𝑒𝑙𝑛

(𝑚𝑓

𝑚𝑖

)(2.34)

where 𝑉𝑒 is the exhaust velocity as it leaves the nozzle, 𝑚𝑓 is the final mass of the

satellite after the propellant is ejected and 𝑚𝑖 is the initial mass of the satellite including

the propellant. Though most of the propellant is not ejected from the satellite, effective

exhaust velocity, effective specific impulse and the effective ΔV of the proof of concept

design will be calculated in 𝐶ℎ𝑎𝑝𝑡𝑒𝑟 4.

47


0.7 0.75 0.8 0.85 0.9 0.95 1

Dry Mass Fraction

0

50

100

150

200

250

300

350

400

450

De

lta V

(m

/s)

X: 0.8789Y: 149.4

Figure 2.5: ΔV performance versus the dry mass fraction for an 𝐼𝑠𝑝 of 118 𝑠

The design target performance values and the design parameters are summarised in Table

2.3, with the nozzle half-angle, 𝜃𝑐𝑛, chosen in compliance with SSTL standard for the

design of a nozzle for a thrust efficiency of 98.5% from 𝜆 = 12 (1 + cos 𝜃𝑐𝑛) [131].

Table 2.3: Hybrid propulsion system design parameters and performance

𝐹 Thrust, 𝑁 0.29

𝑎𝑜 Acoustic velocity, 𝑚/𝑠 520.18

𝑐* Characteristic velocity, 𝑚/𝑠 673.166

𝐼𝑠𝑝 The specific impulse, 𝑠 118

ΔV Net velocity change to the spacecraft, 𝑚/𝑠 150

𝑀𝑒 Exhaust Mach number 5.03

𝑝𝑒 Exit pressure, 𝑏𝑎𝑟 0.0052

Propellant mass flow rate, 𝑘𝑔/𝑠 0.00025

𝐴𝑡 Nozzle throat cross-sectional area, 𝑚𝑚2 0.42544

𝐷𝑡 Nozzle throat cross-sectional diameter, 𝑚𝑚 0.74

𝑒 Nozzle expansion ratio 40

𝜃𝑐𝑛 Nozzle half-angle, ∘ 14

𝐴𝑒 Nozzle exit area, 𝑚𝑚2 17.0176

𝐷𝑒 Nozzle exit diameter, 𝑚𝑚 4.65

48


2.4.1 Thruster Design

An experimental proof of concept to demonstrate the outlined objectives in Chapter 1

has been developed. There are two basic elements of the design: the reaction chamber

and nozzle. The thruster was designed and mounted on a moving plate of a thrust

balance for experimental thrust measurements. The components of the thruster are

described in this section.

2.4.1.1 Reaction Chamber

The chemical reaction of the propellants takes place in the reaction chamber. The

reaction chamber helps to convert the released chemical enthalpy and the random motion

of the exhaust gases into directed kinetic energy through the nozzle. Traditionally the

design of the thrust chamber is based on requirements that define its major parameters.

These requirements include the defined mission of the spacecraft, area of operation,

repeatability of the chemical process, reliability, mass and size of the thruster. At this

early stage of proof of concept testing of the model, our emphasis is on the repeatability

of the propulsion system and its ability to fit into the mass and volume of the spacecraft.

Since the chemical reaction in the chamber is non-combusting, there was no restriction

on its length and area, except to suit a CubeSat restricted volume. Hence the scaling

effect [141] of a characteristic chamber length for complete combustion in a conventional

chamber [142] was not considered in this design. The major requirement in the design is

for the chamber to contain as much propellant as needed for a 1𝑈 CubeSat as calculated

in section 2.3, and the chamber wall to withstand the generated pressure inside the

reaction volume. The reaction chamber was made from a 316 stainless steel and was

measured 42𝑚𝑚 × 30𝑚𝑚 in height and diameter respectively and a wall thickness of

1𝑚𝑚, just enough to take about 6 𝑔 fuel for the instance. The design of the reaction

chamber is shown in Figure 2.6, with the top images showing the inside of the chamber

while the bottom image shows the assembled part. The detailed manufacturing drawings

are provided in 𝐴𝑝𝑝𝑒𝑛𝑑𝑖𝑥 𝐴.

49


AlldLenghtsdindmmAnglesd±X°dDistanced±dX SIZEjA4

PartdNamejdExpNozzlePartXXGdftScalejdXdjdX

Materialj

PartsdRequiredj

SurreydSpacedCentred)SSCUUniversitydofdSurreyBdBAdBuilding

GuildfordBdSurreyBdGUHdOXH

H7/7k/H7X7AhmeddOzomatadDavid

oa77XXO@surreyGacGuk 7O7 77 X

SteelXdX

ProjectdSheetjd Xdo

Oxidiserflow inlet Reaction

chamber

Temperature andpressure sensors

port

Nozzleport

Figure 2.6: Design model of the reaction chamber

The two protrusions on the side of the chamber are for the thermocouple and pressure

sensors to read the chamber temperature and pressure respectively.

2.4.1.2 Nozzle

Chemical propulsion systems use nozzles to direct and accelerate the chemical reaction

products to maximize the exhaust velocity at the exit. The nozzle has two basic sections:

converging and diverging, with an adjoining part known as the throat. The exhaust

from the reaction chamber is first converged into a subsonic flow to the throat. The flow

becomes choked at the throat with a unity Mach number. Afterwards, the flow begins to

expand isentropically at the diverging section to a supersonic flow. This is made possible

by the generated chamber pressure that pushes the exhaust products through the nozzle.

This pressure decreased through both the converging and diverging sections of the nozzle

to the exit. The thrust generated by the nozzle is determined by the exit velocity of the

exhaust gas, the propellant mass flow rate, the ambient pressures and the nozzle area,

as seen in Equation 2.29. This concept was implemented in this research by designing

a simple converging-diverging nozzle with a throat area based on the thermodynamic

analysis of Section 2.3 using the listed parameters in Table 2.3. Two different nozzle

expansion ratios were adopted in this study to compare the performance characteristics

of the propulsion system. The nozzle was constructed from an adapted swagelok cap

50


and plug, and the design of the converging section of the nozzle is shown in Figure 2.7.

The detailed drawing of the nozzle is provided in 𝐴𝑝𝑝𝑒𝑛𝑑𝑖𝑥 𝐴

14o

Figure 2.7: Swagelok cap and plug [143] adopted as nozzle

51

Chapter 3

Experimental Setup

3.1 Overview

The section contains the description of an experimental lab setup to characterise the

propulsion system. This includes the description of the oxidiser feed system, data

acquisition system, vacuum facilities and the complete experimental setup.

3.2 Oxidizer Feed System

The propellant feed system determines how much of the sodium hydroxide + water

oxidiser is delivered per unit time to the reaction chamber containing the aluminium wool.

It consists of a tank to store the propellant, a feed mechanism that pushes the propellant

from the tank into the reaction chamber, piping, and a control mechanism to initiate

and regulate the propellant flow rate [144, 71]. Two conventional types of propellant

feed systems can be found in literature. They are the pump fed and the pressure fed

systems. The pump fed system pressurises the propellant into the reaction chamber

using turbopump at relatively high pressures [144]. This type of feed system is used for

short duration and high total impulse applications. The second type of feed system is a

pressure fed system. This is a simple method as it relies on the tank pressure to push

the propellant into the reaction chamber. This system is used for low total impulse and

long duration orbital missions. We adopted the pressure fed system for this propulsion

system for its simplicity, whose schematic is shown in Figure 3.1. It contains a cool gas

52

3.2. Oxidizer Feed System

generator (from Cool Gas Generator Technologies) that provides about 98% of nitrogen

gas with a pressure range of 10 𝑏𝑎𝑟 [145]. The oxidiser tank houses a chemical resistive

membrane, like a polytetrafluoroethylene (PTFE) membrane [146], that envelopes the

oxidiser as shown in Figure 3.1. The generated pressure compresses the membrane and

allows the oxidiser flows into the plenum volume once the isolation valve is opened. For

the proof of concept design, we used the atmospheric pressure to provide about 1 𝑏𝑎𝑟 to

feed the oxidiser into the reaction chamber, as shown in Figure 3.2.

Oxidiser tank

Cool gas generator

Isolation valve 1

Sodium hydroxide resistive membrane

To plenum volume

Figure 3.1: Schematic of a pressure feed system

Oxidiser tank

Solenoid valve

PFA valve

PVC plastic pipe

Plenum volume

Arduino controller

Figure 3.2: Feed system setup

53

3.3. Data Acquisition System

The feed system consists of a tank with an open end, two GEM-SOL Chem-Sol plastic

solenoid valves (see 𝐴𝑝𝑝𝑒𝑛𝑑𝑖𝑥 𝐵 for details), plain clear PVC plastic 1/4 inches pipe and

four PFA needle vales from swagelok (see 𝐴𝑝𝑝𝑒𝑛𝑑𝑖𝑥 for details). The needle vales were

used as a means of regulating the propellant flow into the reaction chamber. A length of

the plastic pipe in-between the solenoid valves was used as plenum volume, holding a

maximum of 6𝑚𝑙 of oxidiser to be injected into the reaction chamber by the operation

of the two solenoid valves. The control is done through a programmed arduino controller

that is attached to the valves and the connection circuit is shown in Figure 3.3.

GND

¼

5-v

ArduinoUNO

D-

Dv

Val-

Valv

R-

Rv

Rp

V-

Vv

T-

TvSW

Keys

V-

Vv

T-e=eTv

Rve=eRp

SW

D-e=eDve

Symbols Meaning:ValuesDiode:INe-¼,,-

Resistor:-kΩ

Microeswitch

NPNe Transistor:e 8Ae -,,Ve HFE:e v,,HpFPineTOFvv,

5eVedc

-veVedc

Earth

R- Resistor:-,kΩ

Val-e=eValv PlasticesolenoidevalveHeporte¼einches

Figure 3.3: Circuit connection of the arduino and the solenoid valves

The arduino board and the program used in this project are provided in𝐴𝑝𝑝𝑒𝑛𝑑𝑖𝑐𝑒𝑠 𝐵

and 𝐶.

3.3 Data Acquisition System

Data generated during the course of this experiment were measured, processed and

stored using a data acquisition system (DAQ). The system consists of temperature and

pressure sensors, and a National Instruments NIDAQ card using a custom LabVIEW

interface.

3.3.1 Temperature and Pressure Sensors

Two types of temperature sensor were used to read the temperature of the thruster. They

are both type K but different in characteristics. A type K insulated thermocouple with a

54


sensitivity of approximately 41𝜇𝑉 /∘C and a temperature range of -40 to 1100∘C, Figure

3.4(a), was used for the inside of the reaction chamber while a K type thermocouple with

a temperature range of -50 to 250∘C, Figure 3.4(b), was used to measure the outside

wall temperature of the thruster. These thermocouples were selected because they are

rugged, flexible, and suitable for vacuum applications [147].

(a) (b)

Figure 3.4: Thermocouples: (a) Insulated thermocouple and (b) Fine wire thermocouple

The pressure transducer used in the experiment is shown in Figure 3.5. It is a rugged

PXM309 pressure series with a range of 0 − 70 𝑏𝑎𝑟 at an output voltage of 0 to 10 𝑉 𝑑𝑐

from Omega Engineering Ltd [148].

Figure 3.5: Pressure transducer

3.3.2 DAQ Measurement Hardware

DAQ measurement hardware acts as the interface between the signals from the sensors

and the lab computer. It converts the analogue signals from the sensors into digital format

for the computer. The hardware is made up of a National Instrument (NI) PCI-6221 card,

a terminal block and a shielded cable [149]. Figure 3.6 shows the hardware components

and their connections. Multiple signals from the transducers are routed through the

shielded cable to the PCI card, which is slotted into the computer.

55


To PCI slot

From transducers

PCI-6221 card

Terminal block

Shielded cable

Figure 3.6: DAQ measurement hardware and connections

3.3.3 LabVIEW Software

A custom LabVIEW interface was programmed to monitor, control and automate the

data acquisition process. The block diagram of the LabVIEW program is shown in

Figure 3.7.

56

3.4. Vacuum Facilities and Thrust Balance

Figure 3.7: Block diagram of the LabVIEW program used to control and acquire datafrom sensors

3.4 Vacuum Facilities and Thrust Balance

The propulsion lab of the Surrey Space Centre at the University of Surrey has four

vacuum facilities for high vacuum testing of propulsion systems. These include 𝐷𝑖𝑛𝑘𝑜,

𝐻𝑒𝑟𝑚𝑒𝑠, 𝑃𝑒𝑔𝑎𝑠𝑢𝑠 and 𝐷𝑎𝑒𝑑𝑎𝑙𝑢𝑠 vacuum facilities. The vacuum facilities, which contain

equipment that enables the testing and analysis of propulsion systems, provide adequate

conditions on earth to simulate the space environment and to validate the performance

characteristics of propulsion devices before they are flight ready.

3.4.1 The Pagasus Vacuum Chamber

Th 𝑃𝑎𝑔𝑎𝑠𝑢𝑠 chamber [150] is a horizontal stainless steel cylinder of 2𝑚 diameter and

1.5𝑚 length resulting in a total chamber volume of approximately 4.7𝑚3. The chamber

is connected to a rotary roughing pump and a turbo molecular pump with a combined

pumping speed of approximately 1700 𝑙/𝑠, which remove air, moisture and other gases

from the chamber. On the body of the chamber are multiple ports that are covered

57


with vacuum flanges to allow for the passage of instruments, like sensors, electrical and

gas feeds, into the chamber for a complete experimental set up. The ports also provide

optical access to an experiment in the chamber. The chamber generates a base pressure

of about 0.1𝑚𝑃𝑎 and an operating pressure of 10-100𝑚𝑃𝑎 depending on the propellant

flow rate. The automating pump down and venting sequences of the chamber, like

closing and opening of gate vales, are done through a dedicated desktop computer, which

also provides a means of reading the chamber pressure. Figure 3.8 shows the 𝑃𝑒𝑔𝑎𝑠𝑢𝑠

vacuum chamber that was used in these experiments.

Figure 3.8: 𝑃𝑒𝑔𝑎𝑠𝑢𝑠 vacuum chamber

3.4.2 Thrust Balance Arrangement

The Surrey Space Centre has a pendulum type thrust balance where the thrust produced

by the propulsion system can be evaluated. The thrust is measured by the displacement of

the pendulum under the action of an applied force and achieves a measurement accuracy

of approximately 0.1𝑚𝑁 [151]. The thruster is attached to a flat platform, a moving

plate, and suspended from the thrust stand by an arrangement of flexures and struts,

which allows the thruster to move freely in a horizontal direction as shown schematically

58


in Figure 3.9.

Laser displacement sensor

Moving plate

Electric motor

Calibration weight

Thruster

Struts

Flexures

Laser target

Figure 3.9: Schematic of the thrust stand

The movement of the thruster when displaced is measured by a commercial laser displace-

ment sensor (Micro-Epsilon model ILD 1700-50) that is mounted to a fixed plate of the

thrust stand and aligned with a straight ceramic target that is mounted to the moving

plate, see Figure 3.9. The displaced distance between the target and the laser is linear

proportional to the thrust value of the thruster; the correlation factor is determined

through 𝑖𝑛 𝑠𝑖𝑡𝑢 calibration. The laser operates over a 50𝑚𝑚 range and a manufacturer

reported linearity of 40𝜇𝑚 with a resolution of 3𝜇𝑚 at 2.5 𝑘𝐻𝑧 (See 𝐴𝑝𝑝𝑒𝑛𝑑𝑖𝑥 for detail)

[152].

3.4.3 Calibration and Data Analysis

An 𝑖𝑛 𝑠𝑖𝑡𝑢 calibration of the stand is done each time the experiment is performed with

a known mass of 10 𝑔 attached to an electric motor at a fixed distance from the stand

frame with an inextensible thread. The electric motor used during calibration is a DC

stepper motor from RS Components Limited [153] and shown in Figure 3.10.

59


Figure 3.10: DC stepper motor from RS Components Limited

The motor rotates forward and reverses depending on commands received from an arduino

controller interfaced to the desktop computer. The geometry of the displacement of

the weight with respect to the moving plate of the thrust stand are shown in Figure

3.11.

θ

22 sls

mg−

=

zF

(a) (b)

h

s

mg

l

Fh

l

Figure 3.11: Geometry Of Thrust Calibration and Thrust Stand Sepup with StepperMotor

The calibration weight is moved in a step-wise sequence consisting of 6 forward steps

and 6 backward steps of equal length. The horizontal force applied to the thrust stand,

when the calibration weight is moved, is determined by an equation given by [154]

𝐹ℎ =𝑚𝑐𝑔𝑜𝑠√𝑙2 − 𝑠2

, (3.1)

where 𝐹ℎ is the horizontal force, 𝑚𝑐 is the mass of the calibration attached to the stand,

60


𝑠 is the horizontal displacement distance of the mass from the stand, and 𝑙 is the length

of the pendulum thread suspending the mass. The thrust stand has an accuracy of

±0.1𝑚𝑁 [155].

The plots of the moving plate in response to the applied force as a function of time is

shown in Figure 3.12(a). The stable position of the plate before the action of the electric

motor is indicated by about 21.9𝑚𝑚 mark on the vertical axis. At this point there is

no weight on the horizontal thread attached to the motor. The red signal represents

the addition of weight in the horizontal axis by the displacement of the moving plate

in forward and backward direction. The associated noise in the signal was filtered out

using a butterworth low pass filter, which is shown in blue line in the same figure. The

horizontal green line shows a repetition of the corresponding positions of the moving

plate when moving backward. Figure 3.12(b) shows the linear relationship of the thrust

balance to the applied force. Though six positions were considered for the thrust balance

calibration, the 50𝑚𝑚 range of the laser displacement sensor could accommodate a

higher force than applied. The gradient of the relationship of a particular calibration

gives a calibration constant of about 2.541𝑚𝑚/𝑁 or 0.39355𝑁/𝑚𝑚, and it gives the

corresponding relationship between the applied force and the deflection of the moving

plate.

0 10 20 30 40 50 60

21.75

21.8

21.85

21.9

21.95

Time (s)

Dis

pla

cem

en

t (m

m)

0.01 0.02 0.03 0.04 0.05 0.06 0.0721.75

21.8

21.85

21.9

Force (N)

Dis

pla

cem

en

t (m

m)

(a) (b)

Figure 3.12: Responses of Thrust Calibration

61

3.5. Complete Experimental Setup

3.5 Complete Experimental Setup

This section describes the integrated experimental setup. It also shows the flow path

of propellant from the oxidiser tank into the reaction chamber already containing fuel.

Figure 3.13 shows the schematic of the complete experimental setup.

The laser displacement sensor of Section 3.4.2, electric stepper motor and the cali-

bration weight of Section 3.4.3 are attached to the thrust stand and placed in the vacuum

chamber, with all the electrical cables interfaced through feedthroughs and connected

to the computer. At the start of each experiment, the thruster is pre-loaded with a

precise mass of aluminium wool fuel. A controlled volume of oxidiser, by the operation

of the two plenum valves, is released from the oxidiser tank into the reaction chamber to

commence the reaction. The rate of temperature and pressure change within the reaction

volume are recorded using the LabVIEW software program of Section 3.3.3 through their

respective sensors. The resulted thrust force produced by the device is sensed by the

laser sensor and recorded by the computer.

62

3.5. Complete Experimental Setup

Computer

Terminal block for D-sub 37 pin cable

USB-COMi-M

Arduino card2

Arduino card1

Desktopcomputer

Vacuum chamber

Control volume

Valve1 Valve2

Laser sensor signal

Temperature and pressure sensors’ signalsNI-DI shielded cable

12V/2A

12V/0.06A

Oxidiser

Figure 3.13: Schematic of the complete experimental setup

63

Chapter 4

Results and Discussion

4.1 Overview

This section focuses on the experimental results obtained during this thesis and their

analysis for the evaluation of the thruster. The experimental results are also compared

with theoretical values as well as the state-of-the-art technology in this section.

4.2 Stages of Experiments

The first stage of the experiment was to evaluate the reaction under ambient conditions

between aluminium wool and sodium hydroxide solution with varying molarity. In later

stages of the experiment, the reaction volume was situated within the vacuum chamber,

and pressure was allowed to build up as a function of time using a restricted orifice.

4.2.1 Reaction Chemistry of the Propellants at Ambient Conditions

Molality (also known as molar concentration) of a solution is the number of moles of a

solute present in 1 𝑘𝑔 of solvent to form the solution. It is measured in 𝑚𝑜𝑙/𝑘𝑔. Molality

is preferred in this experiment because it involves temperature changes and molality is

based on mass rather than volume, which increases when its temperature increase. That

is, heating makes the molarity of a solution go down. In this experiment, the molality of

sodium hydroxide that will react with aluminium wool to give the highest temperature

64

4.2. Stages of Experiments

rise was determined by adding the same amount of the aluminium wool to different

beakers containing the same amount of solvent (water) with different moles of sodium

hydroxide under ambient conditions. That is, different molality of solution of sodium

hydroxide. Two parameters were used to evaluate the reaction: reaction time and rise in

temperature. The schematic of the experiment set up is shown in Figure 4.1. 0.001 𝑘𝑔 of

aluminium wool was added to each of the beakers with different molality, ranging from

1.56𝑚𝑜𝑙/𝑘𝑔 to 25.00𝑚𝑜𝑙/𝑘𝑔. Four K-type thermocouples were attached to the back of

the beakers with a crypton tape to relatively measure the temperature rise in each baker.

The summary of the experimental result is shown in Table 4.1.

Thermocouple

Aluminium wool Crypton tapeSodium hydroxide solution

Beaker

1 2 3 4

Kapton tape

Figure 4.1: Experimental setup for sodium hydroxide concentration on aluminium-waterreaction

Table 4.1: Experiment data for sodium hydroxide molality

Sample 1 2 3 4

Mass of water (kg) 0.01 0.01 0.01 0.01

Moles of sodium hydroxide

(mol)

0.02 0.06 0.12 0.25

Mass of aluminium (kg) 0.001 0.001 0.001 0.001

Molality of sodium hydroxide

(mol/kg)1.56 6.25 12.50 25.00

Figure 4.2 shows the temperature rise against time for the various aluminium and

water-sodium hydroxide reactions. Molar concentrations of 6.25𝑚𝑜𝑙/𝑘𝑔 has the highest

temperature rise but its induction time is about 100 𝑠, which is longer than both of

molalities 12.50𝑚𝑜𝑙/𝑘𝑔 and 25.00𝑚𝑜𝑙/𝑘𝑔 whose iduction time is about 30 𝑠. Though

65


the induction time of both molalities 12.50 𝑚𝑜𝑙/𝑘𝑔 and 25.00𝑚𝑜𝑙/𝑘𝑔 is small, molality

12.50𝑚𝑜𝑙/𝑘𝑔 has higher temperature rise than molality 25.00𝑚𝑜𝑙/𝑘𝑔. Also molality

12.50𝑚𝑜𝑙/𝑘𝑔 has the second highest temperature rise among the considered molalites.

Hence, sodium hydroxide-water mixture of molar concentration of 12.50𝑚𝑜𝑙/𝑘𝑔 was

chosen as the propulsion oxidiser molality in this research. However, since the shortest

induction time from these molalites is longer than the response required for precision

orbital control, the propose propulsion system is most suited for gross orbital change

and de-orbiting.

0 200 400 600 800 1000 1200 1400 160020

30

40

50

60

70

80

90

100

Time (s)

Tem

p. (°

C)

Temp1 = NaOH (0.625g)

Temp2 = NaOH (2.5g)

Temp3 = NaOH (5g)

Temp4 = NaOH (10g)

Temp5 = Ref Temp.

Figure 4.2: Effect of sodium hydroxide molality on aluminium-water reaction

4.2.2 Temperature and Pressure Rise in a control volume under Vac-

uum Conditions

The molar concentration of 12.50𝑚𝑜𝑙/𝑘𝑔 from the previous experiments was used in this

section to determine both temperature and pressure rises of the propellant combinations

in a reaction chamber with an orifice area of 0.425𝑚𝑚2 under a vacuum conditions. This

experiment was conducted in a vacuum chamber with a backing pressure of 0.0002 𝑏𝑎𝑟.

The schematic of the experimental set up is shown in Figure 4.3. In this experiment, the

fuel mass was systematically increased from 0.0005 𝑘𝑔 to 0.007 𝑘𝑔 while the oxidiser mass

was repeatedly kept the same, at 0.0038 𝑘𝑔, though with the same molar concentration.

In each experiment, water of 0.0025 𝑘𝑔 was pre-mixed with 0.003125𝑚𝑜𝑙 of sodium

hydroxide to give the 12.50𝑚𝑜𝑙/𝑘𝑔 molar concentration of the oxidiser. The mixture

66


was shaken to avoid the solidification of the alkaline solution, allowed it to cool down to

dissipate the generated heat before injecting the mixture into the reaction chamber.

Thermocouple

Flow restrictor

Computer

Reaction Chamber

Restrictor Valve

Solid Fuel

Oxidiser Injector

Vacuum Chamber

Feedthrough

Feedthrough

Pressure relief valve

P Pressure gauge

Figure 4.3: Schematic of the initial lab setup

Each experiment was repeated at least twice to confirm the results. The summary of the

experiment is shown in Table 4.2.

Table 4.2: Temperature and pressure rise in vacuum condition

Experiments 1 2 3 4 5

Mass/Concentration of

oxidiser (kg)/(mol/kg)0.00375/12.50 0.00375/12.50 0.00375/12.50 0.00375/12.50 0.00375/12.50

Mass of fuel (kg) 0.0005 0.001 0.003 0.005 0.007

Change in temperature

(∘C)19.00 22.00 40.00 94.00 88.00

Change in pressure (bar) 0.06 0.12 0.37 0.83 0.47

67


0 1 2 3 4 5 6 7

Mass of fuel (g)

0

20

40

60

80

100

Tem

p. (

°C)

(a) Temperature rise in the reaction chamber

0 1 2 3 4 5 6 7

Mass of fuel (g)

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

Pre

ss. (b

ar)

(b) Pressure rise in the reaction chamber

Figure 4.4: Changes in temperature and pressure in the reaction for the same mass andconcentration of oxidiser but different mass of fuel

Figure 4.4 shows the variations in the temperature and pressure in the different combi-

nations of the propellants. The highest temperature and pressure changes of 94 ∘C and

0.83 𝑏𝑎𝑟 respectively were recorded for the reaction between fuel of 0.005 𝑘𝑔 and oxidiser

of 0.00375𝑘𝑔. These optimum values occur at a fuel to oxidiser mass ratio of about 1:1.

Further increase in fuel mass did not represent a progressive increase in temperature and

pressure because oxidiser as the limiting reactant of the reaction could not access all the

fuel in the reaction, which was covered by the aluminium hydroxide product that was

formed.

4.2.3 Effect of Oxidiser Concentration on Thruster Characterisation

This set of experiments was performed to demonstrate the effect of oxidiser concentration

on average thrust level performance of the propulsion system. The fuel and oxidiser

masses were kept the same in all the experiments while the concentration of the oxidiser

was varied from 4.17𝑚𝑜𝑙/𝑘𝑔 to 20.83𝑚𝑜𝑙/𝑘𝑔 in five sets of experiments as shown in

Table 4.3. The injection of the oxidiser into the reaction chamber was done at about

10 𝑠 into each of the experiments. This is indicated in the thrust versus time graphs

of Figure 4.5 by spikes, which is due to release of trapped air in the feed system into

the reaction chamber. There are high induction times and non-smooth thrust responses

for concentrations 4.17𝑚𝑜𝑙/𝑘𝑔, 8.33𝑚𝑜𝑙/𝑘𝑔 and 20.83𝑚𝑜𝑙/𝑘𝑔 when compare to that of

concentrations 12.50𝑚𝑜𝑙/𝑘𝑔 and 16.67𝑚𝑜𝑙/𝑘𝑔. However, concentration 12.50𝑚𝑜𝑙/𝑘𝑔

68


has the highest thrust response with minimised spikes.

Table 4.3: Data for effect of oxidiser concentration on thrust level

Exp. Mass of Mass of Oxidiser molality Propellant mass after reaction Temperatur Average

fuel oxidiser 𝐻2𝑂 𝑁𝑎𝑂𝐻 Remn’g mass Exht’d mass thrust

(kg) (kg) (kg) (mol) (mol./kg) (kg) (kg) 𝑜C (N)

1 0.004 0.004 0.006 0.025 4.17 0.0058 0.0022 47 0.0017

2 0.004 0.004 0.006 0.050 8.33 0.0056 0.0024 64 0.0023

3 0.004 0.004 0.006 0.075 12.50 0.0059 0.0021 73 0.0045

4 0.004 0.004 0.006 0.100 16.67 0.0061 0.0019 69 0.0034

5 0.004 0.004 0.006 0.125 20.83 0.0062 0.0018 45 0.0015

69


0 50 100 150 200 250 300 350 400

Time (s)

-2

0

2

4

6

8

10

Thr

ust(

N)

# 10-3

(a) Thrust performance for oxidiser concentration4.17𝑚𝑜𝑙/𝑘𝑔

0 50 100 150 200 250 300 350 400

Time (s)

-2

0

2

4

6

8

10

12

14

Thr

ust(

N)

# 10-3

(b) Thrust performance for oxidiser concentration8.33𝑚𝑜𝑙/𝑘𝑔

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

Th

rust

(N)

(c) Thrust performance for oxidiser concentration12.50𝑚𝑜𝑙/𝑘𝑔

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

Th

rust

(N)

(d) Thrust performance for oxidiser concentration16.67𝑚𝑜𝑙/𝑘𝑔

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

0.03

Th

rust

(N)

(e) Thrust performance for oxidiser concentration20.83𝑚𝑜𝑙/𝑘𝑔

Figure 4.5: Thrust reponses for different oxidiser concentration

This is also evident in temperature versus time responses of Figure 4.6, where concentra-

tion 12.50𝑚𝑜𝑙/𝑘𝑔 recorded the better and the highest temperature which is in agreement

70


with section 4.2.1 where 12.50𝑚𝑜𝑙/𝑘𝑔 produced the highest temperature and pressure

responses.

0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

80

Te

mp

. (°

C)

(a) Temperature performance for oxidiser concen-tration 4.17𝑚𝑜𝑙/𝑘𝑔

0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

Tem

p. (

°C)

(b) Temperature performance for oxidiser con-centration 8.33𝑚𝑜𝑙/𝑘𝑔

0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

80

90

100T

emp.

(°C

)

(c) Temperature performance for oxidiser concen-tration 12.50𝑚𝑜𝑙/𝑘𝑔

0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

80

90

100

Tem

p. (°

C)

(d) Temperature performance for oxidiser con-centration 16.67𝑚𝑜𝑙/𝑘𝑔

0 100 200 300 400

Time (s)

20

30

40

50

60

70

80

90

Tem

p. (

°C)

(e) Temperature performance for oxidiser concen-tration 20.83𝑚𝑜𝑙/𝑘𝑔

Figure 4.6: Temperature reponses for different oxidiser concentration

71


The same oxidiser concentration gives the highest average thrust values in this experiment

as shown in Figure 4.7.

4.17 8.33 12.50 16.67 20.83

Oxidiser Concentration (mol./kg)

1

1.5

2

2.5

3

3.5

4

4.5

5

Ave

rage

Thr

ust (

N)

# 10-3

Figure 4.7: Average thrust performance response to oxidiser concentration

4.2.4 Fuel/Oxidiser Ratio on Thruster Performance

The goal of the section is to determine the effect of fuel to oxidiser ratio on thrust

performance of the propulsion system with oxidiser concentration of 12.50𝑚𝑜𝑙/𝑘𝑔 in

all cases. Four different propellant ratios were considered for this exercise as shown

in Table 4.4. 0.003 𝑘𝑔 of fuel was first placed in the reaction chamber and 0.009 𝑘𝑔 of

oxidiser injected into it, making fuel to oxidiser ratio of 1:3. The mass of the oxidiser

was systematically reduced to vary the ratio in the successive experiments. However, the

highest mass of the fuel used in the experiment (0.006𝑘𝑔) was limited by the volume of

the reaction chamber. Before each experiment, the vacuum chamber is pumped down to

a back ground pressure of about 0.0002 𝑏𝑎𝑟 and the injection of the oxidiser was initiated

about 10 𝑠 into each of the experiments.

Table 4.4: Data for fuel/oxidiser effect on one-shot experiment

Exp. Mass of Mass of Ratio Propellant mass after reaction Temperature Average Total Specific

fuel oxidiser Remn’g mass Exht’d mass thrust impulse impulse

(kg) (kg) (kg) (kg) 𝑜C (N) (Ns) (s)

1 0.003 0.009 01:03 0.0061 0.0059 46 0.0036 1.2793 22.22

2 0.003 0.006 01:02 0.0064 0.0026 52 0.0015 0.382 14.86

3 0.003 0.003 01:01 0.0044 0.0016 70 0.0036 0.2578 16.53

4 0.006 0.003 02:01 0.0075 0.0015 117 0.0047 0.6792 44.96

Figure 4.8(a) and Figure 4.8(b) have similar performance and both having irregular

72


responses with longer settling time, though the former has higher thrust value than the

later. They are respectively for propellant ratios 1:3 and 1:2. On the basis of these, it

appears that excessive oxidiser to fuel diminishes the reaction temperature and reaction

rate and results in poor propulsive performance. On the contrary, Figure 4.8(c) and

Figure 4.8(d) of ratios 1:1 and 2:1 respectively have shorter settling time, higher thrust

values and smoother responses with higher propellant efficiency, providing a possibility of

repeatable injections. The spikes at the instance of oxidiser injection are due to release of

trapped air in the feed system. These later propellant ratios also show higher temperature

responses as indicated in the table and shown in Figure 4.9. Hence propellant ratio 2:1

gives the highest average thrust of 0.0047𝑁 and specific impulse of about 45 𝑠 for the

considered combinations.

0 50 100 150 200 250 300 350 400

Time (s)

-5

0

5

10

15

20

Th

rust

(N)

# 10-3

(a) Thrust performance for fuel to oxidiser ratioof 1:3

0 50 100 150 200 250 300 350 400

Time (s)

-2

0

2

4

6

8

Thr

ust(

N)

# 10-3

(b) Thrust performance for fuel to oxidiser ratioof 1:2

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

Thr

ust(

N)

(c) Thrust performance for fuel to oxidiser ratioof 1:1

0 50 100 150 200 250 300 350 400

Time (s)

-0.02

-0.01

0

0.01

0.02

0.03

0.04

Th

rust

(N)

(d) Thrust performance for fuel to oxidiser ratioof 2:1

Figure 4.8: One-shot thrust characterisation of the propulsion system on differentpropellant ratios

73


0 200 400 600 800

Time (s)

20

30

40

50

60

70

80

Tem

p. (

°C)

(a) Temperature performance for propellant ratio1:3

0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

Te

mp

. (°

C)

(b) Temperature performance for propellant ratio1:2

0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

80

90

100

Te

mp

. (°

C)

(c) Temperature performance for propellant ratio1:1

0 200 400 600 800 1000

Time (s)

20

40

60

80

100

120

140

Te

mp

. (°

C)

(d) Temperature performance for propellant ratio2:1

Figure 4.9: Temperature reponses for different propellant ratios

4.2.5 Propellant Mass Effect on Thruster Performance

In this section, the ratio of the propellants remained constant while their masses were

sequentially increased to the limit of the reaction chamber to determine the effect on the

propulsion system. Table 4.5 shows the experimental data, together with the temperature

reading, average thrust level, the total impulse and the specific impulse for each of the

scenarios. Though the pattern of thrust response is similar in all the cases, there is a

relative increase in all the performance parameters considered as the propellants mass

increases. This is evident in the table and in Figure 4.10. The increase in the total

impulse as the mass increases is due to a longer settling time of the thrust responses at

higher masses.

74


Table 4.5: Experimental data for variation in propellant mass at constant ratio

Exp. mass of Mass of Ratio Propellant mass after reaction Temperature Average Total Specific

fuel oxidiser Remn’g mass Exht’d mass thrust impulse impulse

(kg) (kg) (kg) (kg) 𝑜C (N) (Ns) (s)

1 0.003 0.003 01:01 0.0044 0.0016 62 0.0037 0.2578 16.53

2 0.004 0.004 01:01 0.0059 0.0021 67 0.0035 0.5580 27.2157

3 0.005 0.005 01:01 0.0076 0.0024 69 0.0054 0.6860 29.2588

4 0.006 0.006 01:01 0.0086 0.0034 70 0.0036 1.4268 42.6520

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

Thr

ust(

N)

(a) Thrust performance for fuel mass 0.003 𝑘𝑔and oxidiser mass 0.003 𝑘𝑔

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

Th

rust

(N)

(b) Thrust performance for fuel mass 0.004 𝑘𝑔and oxidiser mass 0.004 𝑘𝑔

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

Th

rust

(N)

(c) Thrust performance for fuel mass 0.005 𝑘𝑔and oxidiser mass 0.005 𝑘𝑔

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

Th

rust

(N)

(d) Thrust performance for fuel mass 0.006 𝑘𝑔and oxidiser mass 0.006 𝑘𝑔

Figure 4.10: Thruster characterisation at different propellant mass combinations

However, the temperature responses for all the scenarios are relatively the same because

the propellant combinations are in the same ratio, except that at higher mass more of

the fuel is exposed to more oxidiser resulting in higher temperature. The temperature

responses are shown in Figure 4.11.

75


0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

80

90

100

Te

mp

. (°

C)

(a) Temperature response for fuel mass 0.003𝑘𝑔and oxidiser mass 0.003 𝑘𝑔

0 200 400 600 800 1000

Time (s)

20

40

60

80

100

120

Te

mp

. (°

C)

(b) Temperature response for fuel mass 0.004𝑘𝑔and oxidiser mass 0.004 𝑘𝑔

0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

80

90

Te

mp

. (°

C)

(c) Temperature response for fuel mass 0.005 𝑘𝑔and oxidiser mass 0.005 𝑘𝑔

0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

80

90

100

Tem

p. (°

C)

(d) Temperature response for fuel mass 0.006𝑘𝑔and oxidiser mass 0.006 𝑘𝑔

Figure 4.11: Temperature reponses for different propellant mass combinations

4.2.6 Impact of Repeat Cycles on Thruster Performance

Different fuel/oxidiser ratios are considered in this section to evaluate the feasibility of

multiple injections (repeat cycles) as shown in Table 4.6. Starting with ratio 1:1, only

one cycle is possible (see Figure 4.12(a)). By increasing the fuel to oxidiser ratio we

can achieve a greater number of repeat cycles as shown in Figure 4.12(b) and Figure

4.12(c). The corresponding temperature responses are shown in Figure 4.13. The

non-uniformity in the thrust and temperature levels of the repeat cycles is due to the

aluminium passivation layer removal, which only occurs in the early injections. After

the removal, the peak of the responses is reached as more energy from the exothermic

reaction is released. The responses then reduce almost exponentially as more oxidiser is

injected into the fuel and the remaining fuel is covered by aluminium hydroxide. However,

76


the possibility of repeat cycles as more fuel mass react with less mass of oxidiser, has

demonstrated that this alternative propulsion system can sustain multiple firing before

the fuel bed is depleted.

Table 4.6: Experimental data for scaling effect and repeat cycles

Exp. mass of Mass of Ratio No of Propellant mass after reaction Total

fuel oxidiser injections Remn’g mass Exht’d mass impulse

(kg) (kg) (kg) (g) (Ns)

1 0.003 0.003 01:01 1 0.0044 0.0016 0.2578

2 0.006 0.004 1.5:01 5 0.0170 0.0091 2.4016

3 0.016 0.004 04:01 8 0.028 0.0200 1.0506

0 50 100 150 200 250 300 350 400

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

Thr

ust(

N)

(a) One repeat cycle for propellant ratio 1:1

0 1000 2000 3000 4000 5000 6000 7000 8000 9000

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

Th

rust

(N)

(b) Five repeat cycles for propellant ratio 1.5:1

0 1000 2000 3000 4000 5000 6000 7000 8000 9000

Time (s)

-2

-1

0

1

2

3

4

5

6

7

Th

rust

(N)

# 10-3

(c) Eight repeat cycles for propellant ratio 4:1

Figure 4.12: Scaling of propellant ratio for more repeat cycles

77


0 200 400 600 800 1000

Time (s)

20

30

40

50

60

70

80

90

100

Te

mp

. (°

C)

Reaction Chamber Wall Temperature

(a) Temperature response for fuel mass 0.003𝑘𝑔and oxidiser mass 0.003 𝑘𝑔

0 1000 2000 3000 4000 5000 6000 7000 8000 9000

Time (s)

0

10

20

30

40

50

60

70

80

90

Te

mp

. (°

C)

(b) Temperature response for fuel mass 0.004𝑘𝑔and oxidiser mass 0.004 𝑘𝑔

0 1000 2000 3000 4000 5000 6000 7000 8000 9000

Time (s)

0

10

20

30

40

50

60

70

Te

mp

. (°

C)

(c) Temperature response for fuel mass 0.005 𝑘𝑔and oxidiser mass 0.005 𝑘𝑔

Figure 4.13: Temperature reponses for different propellant mass combinations

4.2.7 Effect of Nozzle throat Diameter on Thruster Performance

Two different nozzle throat diameters of 0.7𝑚𝑚 and 1𝑚𝑚, with the same expansion

ratio, were considered for their effects on the propulsion system. Table 4.7 shows the

experimental outcome. The thrust level performance for 1𝑚𝑚 throat is about 46% higher

and smoother than the 0.7𝑚𝑚 throat as shown in Figure 4.14, and the corresponding

temperature responses are shown in Figure 4.15. Effect of surrounding vibration is

noticeable on the thrust response as it raises the settling point above the zero line.

78


Table 4.7: Experimental data on the effect of different nozzle throat diameter

Exp. Mass of Mass of Nozzle No of Propellant mass after reaction Total

fuel oxidiser diameter injections Remn’g mass Exht’d mass impulse

(kg) (kg) (𝑚𝑚) (kg) (kg) (Ns)

1 0.006 0.004 0.7 5 0.0190 0.0070 0.0016

2 0.006 0.004 1.0 5 0.0170 0.0091 0.0024

0 1000 2000 3000 4000 5000 6000 7000 8000 9000

Time (s)

-5

0

5

10

15

20

Th

rust

(N)

# 10-3

(a) Thrust response for 0.7𝑚𝑚 nozzle throat

0 1000 2000 3000 4000 5000 6000 7000 8000 9000

Time (s)

-0.005

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

Th

rust

(N)

(b) Thrust response for 1𝑚𝑚 nozzle throat

Figure 4.14: Thrust level performance for different nozzle throat diameter

0 2000 4000 6000 8000 10000

Time (s)

20

40

60

80

100

120

Tem

p. (

°C)

(a) Temperature response for nozzle throat diam-eter 0.7𝑚𝑚

0 1000 2000 3000 4000 5000 6000 7000

Time (s)

20

30

40

50

60

70

80

90

Tem

p. (

°C)

(b) Temperature response for nozzle throat diam-eter 1.0𝑚𝑚

Figure 4.15: Temperature reponses for different nozzle throat diameters

79

4.3. Reaction Pattern of the Propulsion System

4.3 Reaction Pattern of the Propulsion System

Data from the multiple injections experiments were analysed in this section for the

chemical reaction pattern of the propulsion system. These data are summarised in Table

4.8. There are three sections of the reaction pattern as observed in Figures 4.16(a),

(b) and (c) for reaction temperature, peak thrust and total impulse of the propulsion

system: these are upward slope, the peak and downward slope of the graph respectively.

The upward slope is the period of aluminium oxide/passivation layer removal, which is

characterised by slow reaction of the propellants at the first injection. The next phase of

the reaction is when more aluminium is already exposed for more reaction giving rise

to higher enthalpy of the reaction. At this point, the reaction is at peak with rapid

formation of aluminium hydroxide. The last phase is characterised by downward slope

towards the termination of the reaction. At this stage, any more injection of the oxidiser

will rapidly increase the dry mass fraction with less exhaust products due to the limited

enthalpies of the reaction.

Table 4.8: Experimental data for the chemical reaction model of the thruster

No of Experiments

repeat Tempt (∘C) Thrust (N) Total Impulse (Ns)

cycle 1 2 3 4 1 2 3 4 1 2 3 4

1 67.20 52.34 59.10 80.30 0.007883 0.02287 0.00221 0.007089 0.0700 0.3475 0.0634 0.2751

2 84.20 103.60 91.23 111.00 0.03499 0.1739 0.007009 0.01625 0.4665 1.5147 0.3017 0.6581

3 80.30 61.90 95.62 85.45 0.03332 0.004782 0.00347 0.007057 0.7574 0.3629 0.3449 0.2616

4 72.82 61.27 76.09 80.62 0.01272 0.007901 0.00125 0.007431 0.6188 0.0974 0.0905 0.3424

5 70.01 72.66 0.009892 0.003102 0.4889 0.0462

80

4.4. Energy Conversion Efficiency of the Propulsion System

1 1.5 2 2.5 3 3.5 4 4.5 5

No of cycles

50

60

70

80

90

100

110

120

Te

mp

. (°

C)

F(6g), O(4g), 1mmF(4g), O(4g), 1mmF(6g), O(3g), 0.7mmF(6g), O(4g), 0.7mm

(a) Temperature response pattern of the reaction

1 1.5 2 2.5 3 3.5 4 4.5 5

No of cycles

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

Th

rust

(N

)


(b) Thrust response pattern of the reaction

1 1.5 2 2.5 3 3.5 4 4.5 5

No of cycles

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

I-to

t(N

s)


(c) Total impulse response pattern of the reaction

Figure 4.16: Reaction pattern of the thruster parameters

4.4 Energy Conversion Efficiency of the Propulsion Sys-

tem

Table 4.9 shows the data of an experiment that was used to determine the energy

utilisation efficiency of the propulsion system. Again, the injection of the oxidiser is

10 𝑠 into the experiment but the reaction was noticed at about 50𝑠. The 40 𝑠 induction

period is due to the passivation layer on the aluminium wool. It is after the removal

of this layer that there was reaction between the fuel and the oxidiser. The amount of

the fuel that reacted in the reaction will be calculated in this section together with the

released energy. The time trends of thrust, temperature and pressure (this is the only

recorded pressure reading in the whole experiment) used in this analysis are shown in

Figure 4.17.

81


Table 4.9: Data for a one-shot experiment

Mass of Mass of No of Propellant mass Exhaust Thrust Specific Total

fuel oxidiser injections Remn’g mass Exht’d mass composition (%) impulse impulse

(kg) (kg) (kg) (kg) 𝐻2(𝑔) 𝐻2𝑂(𝑔) (N) (s) (Ns)

0.006 0.003 1 0.0075 0.0015 1.77 98.23 0.03174 45 0.6792

0 50 100 150 200 250 300 350 400

Time (s)

-0.02

-0.01

0

0.01

0.02

0.03

0.04

Th

rust

(N)

(a) Thrust performance for one-shot experiment

0 50 100 150 200 250 300 350 400

Time (s)

0

20

40

60

80

100

Te

mp

. (°

C)

(b) Temperature response in the reaction cham-ber

0 50 100 150 200 250 300 350 400

Time (s)

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

Pre

ss. (b

ar)

(c) Pressure response in the reaction chamber

Figure 4.17: Thrust, temperature and pressure responses of a one-shot experiment forenergy efficiency analysis

After the reaction, the remaining propellant mass in the reaction chamber represents the

mass of unused aluminium wool, mass of sodium hydroxide and the mass of aluminium

hydroxide. This can be put mathematically as:

𝐴𝑙(𝑢𝑛𝑢𝑠𝑒𝑑) +𝑁𝑎𝑂𝐻 +𝐴𝑙 (𝑂𝐻)3 = 0.0075𝑘𝑔 (4.1)

But the total mass of aluminium is 0.006 𝑘𝑔 and the mass of sodium hydroxide is 0.001 𝑘𝑔.

82


Hence, the mass of the hydroxyl ions is 0.005 𝑘𝑔 ((𝑂𝐻)3 = 0.005 𝑘𝑔). This is equal to

0.009054𝑚𝑜𝑙𝑒𝑠 (molar mass of 3 hydroxyl ions is 51.024 𝑔/𝑚𝑜𝑙𝑒). Using Equation 4.2 as

a reference,

𝐴𝑙(𝑠) + 3𝐻2𝑂(𝑙) −→ 𝐴𝑙(𝑂𝐻)3(𝑎𝑞) + 1.5𝐻2(𝑔) (4.2)

1𝑚𝑜𝑙𝑒 of𝐴𝑙(𝑂𝐻)3 contains 1𝑚𝑜𝑙𝑒 of𝐴𝑙 and 1𝑚𝑜𝑙𝑒 of (𝑂𝐻)3. Therefore, 0.0090154𝑚𝑜𝑙𝑒𝑠

of𝐴𝑙(𝑂𝐻)3 contain 0.0090154𝑚𝑜𝑙𝑒𝑠 of (𝑂𝐻)3 and 0.0090154𝑚𝑜𝑙𝑒𝑠 of𝐴𝑙, and 0.0135231𝑚𝑜𝑙𝑒𝑠

of𝐻2 is produced from 0.0090154𝑚𝑜𝑙𝑒𝑠 of 𝐴𝑙. Hence, the mass of hydrogen gas is 0.0272 𝑔,

which represents about 2% of the exhaust. Also, the moles of water used in the reaction

is 0.027𝑚𝑜𝑙𝑒𝑠, which is 0.486𝑔 of water out of the injected 2𝑔. The remaining 1.5128𝑔

of water, that is about 98% of the exhaust, changed from liquid to vapour phase during

the reaction as shown in Equation 4.4.

The propulsion system was treated as a closed control system, as seen in Figure 4.18,

to analyse the flow of energy in and out of the system. Equation 4.3 represents the

conservation of energy of the system assuming adiabatic conditions.

ih wchem hh outhmv 2

21

systemClosed

+-

Figure 4.18: Illustration of energy conversion efficiency

ℎ𝑖 + ℎ𝑐ℎ𝑒𝑚 = ℎ𝑤 +1

2𝑚𝑣2 + ℎ𝑜𝑢𝑡, (4.3)

where ℎ𝑖 is the enthalpy of the liquid oxidizer injected into the system, ℎ𝑐ℎ𝑒𝑚 is the

chemical enthalpy of reaction, ℎ𝑤 is the enthalpy component responsible for water phase

change, 12𝑚𝑣2 is the component of kinetic energy for useful thrust, and ℎ𝑜𝑢𝑡 is the

enthalpy of the outgoing stream. There is only one source of non-chemical energy into the

system of Figure 4.18, which is mainly from the injected oxidizer represented by ℎ𝑖. We

assumed the specific enthalpy of water at 23∘ for the oxidiser. Hence, ℎ𝑖 is 1.7395 𝑘𝐽/𝑚𝑜𝑙

[156]. But from Table 4.9, mass of water is 2 𝑔 and sodium hydroxide is 1 𝑔. This gives

83


the moles of injected water to be 0.111𝑚𝑜𝑙𝑒𝑠, and the enthalpy of the oxidiser is 0.1931 𝑘𝐽 .

The specific enthalpy in the chamber is due to the formation of strong chemical bonds

from the chemical reaction which result in the release of heat energy. This liberated

heat has two components: standard enthalpy of formation and a temperature-dependent

change of enthalpy of the reaction as stated in Equations 2.5, 2.6, 2.8. Data from Table

2.2 together with Table 4.10, which shows the Δℎ∘𝑓 values of the propellants are used in

this analysis. The values are obtained from 𝐶𝑅𝐶 𝐻𝑎𝑛𝑑𝑏𝑜𝑜𝑘 𝑜𝑓 𝐶ℎ𝑒𝑚𝑖𝑠𝑡𝑟𝑦 𝑎𝑛𝑑 𝑃ℎ𝑦𝑠𝑖𝑐𝑠

[120].

Table 4.10: Thermodynamic properties of the propellants

Substances Molar mass Enthalpy of formation Specific heat (𝑐𝑝)

kg/mol kJ/mol kJ/mol.K

𝐴𝑙(𝑠) 0.0269815 0 0.0242

𝐻2(𝑔) 0.0020159 0 0.028868

𝐻2𝑂(𝑙) 0.01801528 -285.8 -

𝐻2𝑂(𝑔) 0.01801528 -241.8 0.036031

𝐴𝑙(𝑂𝐻)3(𝑎𝑞) 0.0780036 -1276 -0.128706

The temperature limits, 𝑇1 and 𝑇2, of the second term in Equation 2.5 are based on

the lowest and highest points on the temperature response of Figure 4.17(b). That is

𝑇1 is 23∘ (296.15𝐾) and 𝑇2 is 98∘ (371.15𝐾). Now, with reference to the chemical

reaction of Equation 2.4 and using the thermodynamics Equations 2.5, 2.6 and 2.8,

and data from Tables 2.2 and 4.10, the enthalpy of formation is 418.6 𝑘𝐽/𝑚𝑜𝑙 and the

temperature-dependent term is 3.5471𝑘𝐽/𝑚𝑜𝑙. Hence, the released heat in the reaction

chamber is found to be 415.0529 𝑘𝐽 per mole of aluminium. But only 0.0090154𝑚𝑜𝑙𝑒𝑠 of

aluminium was used in the reaction, giving the total enthalpy of reaction to be 3.7419 𝑘𝐽 .

This is in agreement with the calculated enthalpy of reaction in Section 2.2, where

415.0529 𝑘𝐽 of heat per mole of aluminium was released within a temperature range of

77∘. Therefore, the total enthalpy into the closed system is 3.935 𝑘𝐽 . A portion of this,

ℎ𝑤, is used by a fraction of water component of the oxidiser to change from liquid to gas,

as shown in Equation 4.4.

𝐻2𝑂(𝑙) = 𝐻2𝑂(𝑔) + ℎ𝑤, (4.4)

84


where ℎ𝑤 is the specific enthalpy of water at the pressure and temperature at the instance

of conversion. Data from Figures 4.17(b) and (c) were plotted on a 𝑝 − ℎ diagram

to calculate the specific enthalpy as shown in Figure 4.19. The 𝑆𝑡𝑎𝑟𝑡 𝑡𝑖𝑚𝑒 and the

𝑆𝑡𝑜𝑝 𝑡𝑖𝑚𝑒 on the graph represent the start of the reaction and the stop of the reaction

from the temperature and pressure graphs.

0 500 1000 1500 2000 2500 3000 3500 4000

Enthalpy (kJ/kg)

10-3

10-2

10-1

100

101

Pre

ssu

re (

MP

a)

p-h diagram for water

Start time = 61 s

137.91 kJ/kg 2679.24 kJ/kg

Stop time = 90 s

10 'C

40 'C

60 'C

200 'C

100 'C

400 'C 500 'C 600 'C

300 'C

Figure 4.19: 𝑝−ℎ diagram of water showing the enthalpy-pressure relation in the reactionchamber. The 𝑝− ℎ diagram was drawn from data obtained from [156]. The blue linerepresents saturated liquid water while the red line represents dry saturated steam. Thedome covers water-steam composition with decreasing water content from left to right.

The specific enthalpy required for the phase change is found to be 2541.33 𝑘𝐽/𝑘𝑔 (sub-

tracting the start point from the stop point). But the mass of water vapour from the

exhaust is 1.51275 𝑔, which represents about 97% of the exhaust. Therefore the energy

used for the water phase conversion is 3.8444 𝑘𝐽 .

Another portion of the generated energy is gained by the reaction exhaust gases that

increased their kinetic energy, 12𝑚𝑣2, resulting in the thrust generation. The kinetic

85


energy can also be expressed as a function of the exhaust mass flow rate by:

𝐾𝐸 =1

2𝑚𝑣2

=

∫1

2𝑣2𝑑𝑡

(4.5)

If we assume a constant exhaust velocity, it implies the total impulse can be expressed

as:

𝐼𝑡𝑜𝑡 =

∫𝑣𝑑𝑡

𝐼𝑡𝑜𝑡 = 𝑚𝑣

(4.6)

Equation 4.5 can therefore be rewritten as

𝐾𝐸 =1

2

𝐼2

𝑚2

∫𝑑𝑡

=1

2

𝐼2

𝑚

(4.7)

Substituting values from Table 4.9, the kinetic energy is calculated to be 0.1498 𝑘𝐽 ,

which represents the enthalpy that got converted into the thrust. This value represents

about 3% of the generated heat in the reaction chamber.

In the above analysis, the sum of the kinetic energy and the energy used by the water to

change phase is more than the available energy in the chamber by 0.0592 𝑘𝐽 . To address

the difference, a plot of the reaction enthalpies versus the percentage of water in the

oxidiser that gets into the reaction chamber is obtained, as shown in Figure 4.20.

86


0 2 4 6 8 10 12 14 16 18 20

5% multiples of injected water into the reaction chamber

0

0.5

1

1.5

2

2.5

3

3.5

4

En

erg

y (k

J) hchem

KE+hw

hw

KE

Figure 4.20: Energy iterations for the percentage of water vapour in the system

The point where the generated enthalpy equals the sum of the kinetic energy and the

energy for phase change is represented by the dash-dot line in the figure.

4.4.1 Chemical Analysis of the Residual Propellants

The reaction between the aluminium wool and the mixture of water and sodium hydroxide

is an exothermic reaction that releases hydrogen gas and forms sodium aluminate

as residue. Stoichiometrically, this reaction requires a mixture of 2𝑚𝑜𝑙𝑒𝑠 of sodium

hydroxide in 6𝑚𝑜𝑙𝑒𝑠 of water that will consume 2𝑚𝑜𝑙𝑒𝑠 of aluminium and then release

3𝑚𝑜𝑙𝑒𝑠 of hydrogen gas and form 2𝑚𝑜𝑙𝑒𝑠 of sodium aluminate as shown in Equation

2.2. That is, when 1 𝑘𝑔 of aluminium reacts with 1.6 𝑘𝑔 of sodium hydroxide in 2 𝑘𝑔 of

water, about 0.112 𝑘𝑔 of hydrogen gas will form leaving behind a precipitate of about

4 𝑘𝑔 of sodium aluminate. However in this proof of concept propulsion design and from

Section 4.4, we have reacted only 0.006 𝑘𝑔 of aluminium wool with 0.003 𝑘𝑔 of oxidiser

that contained 0.002 𝑘𝑔 of water and 0.001 𝑘𝑔 of sodium hydroxide. The experimental

calculations revealed that only about 0.00024 𝑘𝑔 of the aluminium wool was consumed

in the reaction. The remaining aluminium wool and the sodium aluminate that formed

constituted to the remaining 0.0075 𝑘𝑔 of propellants in the reaction chamber as shown in

Equation 4.1. These residual propellants were physically examined after the experiment

to see the part of the fuel that was consumed and what actually formed. Figure 4.21(a)

shows the part of the fuel that reacted and the residue of that reaction. Only the part

of fuel closed to the injected oxidiser reacted. A dissection of the fuel also confirms

87


that there was no further reaction inside the fuel due to the formation of the reaction

residue as shown in Figure 4.21(b). The propellant residue was also examined using

(a) Propellant residue after reaction (b) Dissection of propellant residue

Figure 4.21: Physical examination of propellant residue

MicroRaman spectroscopy to determine the microscopic structure of the propellant

remnant at the Advanced Technology Institute (ATI), University of Surrey. A 20𝜇m view

of the residual propellant of two of the experiments shows similarity in the crystalline

structures and the Raman spectra of the samples. However, there is sparse formation of

aluminate in the reaction products due to the size and composition of the aluminium

wool that was used as fuel. Figure 4.22 is the 20𝜇m view of two samples of residual

propellant and Figure 4.23 is the Raman spectra of the propellant residues. There is low

crystallinity in Figure 4.22, which is as a result of the size of the particles of aluminium

in the aluminate. However, the spectral components or wave numbers of the propellant

residue at 1062 𝑐𝑚−1 ±5 𝑐𝑚−1, 3417 𝑐𝑚−1 ±2 𝑐𝑚−1 and 3530 𝑐𝑚−1 ±2 𝑐𝑚−1 confirms

the existence of aluminate in general and the presence of 𝐴𝑙(𝑂𝐻)3 in particular, which

is in agreement with the work of Nsoki [157]. Nonetheless, there were weaker signals of

other species as indicated at wave numbers 253𝑐𝑚−1 and 1067 𝑐𝑚−1, and this confirms

that the aluminium wool is not a pure aluminium but contains some impurities.

88

4.5. Comparison Between Design Target, Theoretical and Prototype Performances

(a) 20𝜇m view of propellant residue sample 1 (b) 20𝜇m view of propellant residue sample 2

Figure 4.22: Microstructure view of two propellant residues

0 500 1000 1500 2000 2500 3000 3500 4000

Wavenumber shift (cm-1)

0

0.5

1

1.5

2

2.5

3

Inte

nsity (

co

un

ts)

# 104

1067

35303617

3419

(a) Raman spectrum of sample 1

0 500 1000 1500 2000 2500 3000 3500 4000

Wavenumber shift (cm-1)

0

2000

4000

6000

8000

10000

12000

14000

16000

Inte

nsity (

co

un

ts)

253

1062

34173530

(b) Raman spectrum of sample 2

Figure 4.23: MicroRamam spectrum analysis of the propellant residues

4.5 Comparison Between Design Target, Theoretical and

Prototype Performances

Some of the assumptions in Section 2.4 for the design target of the propulsion system,

and for the reaction gaseous products, made the model calculations straightforward and

easy, attributing ideal performance to the thruster. For instance, the reaction cham-

ber conditions were put at temperature: 150∘C; and pressure: 400 𝑘𝑃𝑎. We assumed

the gaseous products were ideal, expended isotropically through a choked nozzle and

their percentage combination was water vapour: 95%; and hydrogen gas: 5%. This

lead to 118 𝑠 specific impulse of the propellants that will give about 150𝑚/𝑠 ΔV to a

nanosatellite of 0.88 dry mass ratio. We also assumed a thrust to mass ratio of 0.22𝑁/𝑘𝑔

that will provide a 1.33 𝑘𝑔 CubeSat a 0.2926𝑁 thrust. However, these gases are real

gases that have van der Waals interaction forces among their molecules and they have

89

4.5. Comparison Between Design Target, Theoretical and Prototype Performances

temperature-dependent specific heat capacities, and other attendant properties. Another

factor that led to the lower performance is the position of the feed system from the

reaction chamber. Some portion of the oxidiser could not make through to the reaction

chamber but were adhesively held to the wall of the conducting duck all the way to the

reaction chamber. And at every start of fuel-oxidiser reaction, the build-up pressure

inside the reaction chamber pushed the reaction products through the conducting duck

and the nozzle thereby reducing the performance of the propulsion model. All these

factors will make forgoing performances deviate from reality. For instance, recorded

experimental values for instantaneous temperature and pressure readings in the reaction

chamber were 98∘C and 160 𝑘𝑃𝑎. The average values of these parameters are 71∘C

and 33.38 𝑘𝑃𝑎. Again, the maximum recorded thrust value during the experiment is

0.032𝑁 with an average value of 0.0047𝑁 . Also, the analysis of the experimental result

reveals that the composition of the exhaust is rather 98.23% of water vapour and 1.77%

of hydrogen gas as compared to their assumed composition ratio of 0.95:0.05. The

instantaneous readings are used in this section to calculate the propulsion theoretical

performance while the average values are used to calculate the prototype performance

and the two are compared to the design target of the propulsion system. For example,

a specific impulse of about 45 𝑠 and a ΔV of 57𝑚/𝑠 were recorded for the prototype

performance as compare to 111 𝑠 specific impulse for the theoretical performance.

Summary of the difference between the design target, theoretical performance and

the prototype performance of the design model are shown in Table 4.11. Though

the prototype performance depicts moderate performance of the propulsion system for

nanosatellites, it has provided an alternative approach for CubeSat propulsion that

requires further investigation and research.

90

4.6. Comparison with the State-of-the-Art

Table 4.11: Table of comparison between theory and experimental data

Symbols Meaning Design Theoretical Prototype

target performance performance

𝑇𝑐 Chamber temperature, ∘C 150 98 71

𝑃𝑐 Chamber pressure, 𝑘𝑃𝑎 400 160 33.38

𝐹 Thrust, 𝑁 0.2926 0.032 0.0047

𝛾 Specific heat ratio 1.324 1.33 1.33

𝑎𝑜 Acoustic velocity, 𝑚/𝑠 520.18 487.17 470.18

𝑐* Characteristic velocity, 𝑚/𝑠 673.17 630.45 606.13

Propellant mass flow rate, 𝑘𝑔/𝑠 0.00027 0.00027 0.000011

𝐼𝑠𝑝 The specific impulse, 𝑠 118 111 45

ΔV Net velocity change to the spacecraft, 𝑚/𝑠 150 140 57

4.6 Comparison with the State-of-the-Art

There have been extensive works on the propulsion systems for micro- and nanosatellites

as described in Section 1.2, and including the works of Busek, Marotta, Moog and

Vacco in the development of micropropulsion systems for nanosatellitess. Most of these

systems are in advance stage and are well suiatble for orbital operations that include

attitude control, orbital manoeuvre and rendezvous. However, most of their systems are

electrically based with operating input power in the range of 7.5𝑊 to 50𝑊 , far beyond

the on-board power of a typical 1𝑈 CubeSat. Those with low input power like the

Marotta cold gas micro-thruster and Vacco cold gas propulsion module [4, 158] operate

within 15 𝑏𝑎𝑟 to about 30 𝑏𝑎𝑟 pressure range, which is avoided on CubeSat platform as

the nanosatellites are still carried as a piggy-bag on a launcher for bigger satellites. Other

chemical micropropulsion systems are generally complex for nanosatellites applications,

with heat transfer and scaling related issues as mentioned in Section 1.2. For example,

hydrazine monopropellant systems are hazardous, toxic, carcinogenic and they require

special training for handling procedures and precautions, while the hydrogen peroxide

systems are self-degradable and reduce performance after long storage. However, when

the hybrid propulsion system for CubeSat applications is fully developed, aside that the

propellants are non- hazardous, cheap and storage with no special training for handling,

the on-demand operation pressure builds up gives it a pass to be launched with other

payloads on a launcher. Power demand is about 500𝑚𝑊 only for the operation of valves

91

4.7. Summary of Experimental Findings

in sequence. Also, the storability of the propellants will make it attractive for end-of-life

applications. However, the development of the system is in its early stage with low

technological readiness level and moderate specific impulse.

4.7 Summary of Experimental Findings

The following experiments were conducted and the results analysed for the qualification

of the propulsion model:

∙ The concentration of the oxidiser that give shortest induction period and the highest

reaction temperature and pressure was found to be 12.5𝑚𝑜𝑙/𝑘𝑔. The result was

also confirmed when different concentrations of the oxidiser were used to determine

the thrust performance of the propulsion model.

∙ Different fuel to oxidiser ratios were combined and tested for performance evaluation

of the thruster. First it was discovered that the efficiency of the propellant increased

as more mass of fuel reacted with less mass of oxidiser at a time. And for the

examined ratios, ratio 2:1 had the shortest settling time, highest thrust value of

0.032𝑁 , specific impulse of 45 𝑠 and smoothest thrust response, and also showed

possibility of more repeat cycles.

∙ More repeat cycles was made possible as the mass of fuel increased versus the mass

of oxidiser, and about 5 repeat cycles was demonstrated for the proof of concept

design, when fuel mass of 6 𝑔 reacted with oxidiser mass of 4 𝑔 at a time.

∙ The analysis of the performance responses of the propulsion model revealed that

all the responses followed the same pattern that were classified into three stages:

upward slope, which represented the fuel passivation layer removal, and was

characterised by slow reaction of the propellant at the initial stage; peak level of

the response, which was the point of the reaction when the enthalpy was maximum;

and downward slope, which was toward the termination of the reaction or when

the reaction enthalpy was diminishing to zero.

∙ The energy utilisation of the propulsion system was conducted and it was first

revealed that about 98% of the propulsion exhaust was water vapour and only

about 2% was hydrogen gas. It was also found out that about 3% of the total

92


generated enthalpy was converted into useful thrust, while the remaining percentage

was used by the about 98% of the injected water to change phase from liquid to

gas.

4.7.0.1 Proposed Mechanical Design of the Hybrid Propulsion System

The analysis of performance responses from the experimental results of the propulsion

model reveal that there is about 50 𝑠 induction time before the reaction of the propellants

that also last for 2.5𝑚𝑖𝑛𝑢𝑡𝑒𝑠 with irregular thrust responses. The thrust response is

noticed as soon as the reaction started because the nozzle is opened to the reaction

chamber, which also prevents a build-up of pressure in the chamber and resulted in the

model low performance. This kind of response will essentially limit the model application

to de-orbiting. Solution to this is to include a Lee valve before the nozzle in future

design to enable a build-up of pressure inside the reaction chamber when the once the

propellants reaction starts. The valve will be opened when the operation of the thruster

is needed. In which case, the time response of the thruster will depend on the opening

and closing responses of the valve, making all the thruster performance responses steady

thereby expanding the applications of the propulsion system. The schematic layout of

the propulsion system showing the major components is shown in Figure 4.24.

OxidiserStank

CoolSgasSgenerator

IsolationSvalveS1

IsolationSvalveS2

ReactionSchamber

Nozzle

Fuel

PlenumSvolume

IsolationSvalveS3SodiumShydroxideS

resistiveSmembrane

Figure 4.24: Schematic layout of the propulsion system

The CAD of the proposed hybrid propulsion system shown in Figure 4.25 is designed to

take about one-third of the volume of a 1𝑈 CubeSat. It contains a reaction chamber,

a plenum volume, two oxidiser tanks, two cool gas generators and three Lee extended

performance valves. The system is made up of poly ethyl ethyl ketone (aside from the

off the shelf components like the valves and gas generators) for mass savings and to

provide chemical resistance against the sodium hydroxide oxidiser. Table 4.12 shows the

93


mass prediction of the propulsion system, which is about 30% of a 1𝑈 CubeSat. It is

10×10×3.25 𝑐𝑚3 in volume, which also represents about 30% of the total volume of a

1𝑈 CubeSat.

Micro-coolgas generator

Pipe Lee valve

Systemsupport Platform

Nozzle

Oxidisertank

Reactionchamber

Plenumvolume Oxidiser

tank

(a) Propulsion CAD drawing

10 cm

32.50 cm

10 cm

(b) Propulsion dimensions

Figure 4.25: CAD drawing of hybrid propuldion system for CubeSat applications showingits dimensions in a 1𝑈 CubeSat

The reaction chamber contains the amount of fuel needed for a defined mission, while the

oxidiser tanks have sodium hydroxide resistive membranes that hold the required oxidiser.

A one-bar nitrogen gas generator is included for a blow down of the oxidiser from the

oxidiser tank to the reaction chamber through the plenum volume by the operation of

Lee valves 1 and 2. Valve 3 is only opened once reaction has started and a required

pressure has built up in the reaction chamber. In this prototype model, the oxidiser tank

is designed to hold about 15𝑚𝑙 of oxidiser and the reaction chamber to contain 6 𝑔 of

fuel. The plenum volume of about 3𝑚𝑙 contains a one-shot volume of the oxidiser for a

complete 5 repeat cycles.

94


Table 4.12: Mass estimation of the hybrid propulsion system

Component Mass (g) No of components Total mass (g)

Oxidiser tank 30 1 30

Plenum volume 8 1 8

Reaction chamber and nozzle 10 1 10

Pipe work, connectors and

base bracket

62 1 62

Lee valve 9 3 18

Cool gas generator 2 2 4

Total 132

95

Chapter 5

Conclusions and Future Work

The almost four years journey of this research has taken us though comprehensive litera-

ture about nanosatellites and their micropropulsion systems. The nanosatellites were

initially thought to provide students of higher education institutions expertise in all the

aspect of satellite subsystems. Most of these satellites were launched without propulsion

system except for technological demonstration and were left in their injected orbit until

they spiralled into the atmosphere and de-orbited. The popularity of CubeSats continues

to grow, and non-governmental agencies have become major players in the design and

launching of the satellite due to their low cost. This has necessitated the emergence of

orbital control demonstration on the nanoosatellite platform to expand their capabilities.

We found out that these nanosatellites, especially the CubeSats, were without any form

of propulsion system until about two decades ago when the first nanosatellite to be

flown with propulsion was launched. Since that time a lot of works have been done

in literature in the area of micropropulsion systems for CubeSats. These include the

micro-electrothermal, micro-electrostatic and micro-electromagnetic propulsion systems,

and the various micro-chemical propulsion systems. A comprehensive review in this

area showed that it is difficult to miniaturise well understood and advanced propulsion

systems and still retain their operational advantages, and there is still scope to try a

novel chemical and physical process as the basis of a novel propulsion technology.

In a quest for an alternative chemical process in the design of a novel micropropul-

sion system for CubeSats, we considered water splitting reaction with aluminium wool

fuel using sodium hydroxide as a promoter. The choice of the propellants was based

96

CHAPTER 5. CONCLUSIONS AND FUTURE WORK

on the availability and cost of materials; long storage without degrading; moderate

temperature and exothermic reaction without any thermal control threat to the CubeSat

structure; and high energy density per unit volume for the volume constraint satellite.

Various efforts were identified in literature where the chemical combination was used

for an 𝑖𝑛− 𝑠𝑖𝑡𝑢 generation of hydrogen gas as a source of energy, and as a substitute

for depleting fossil fuels. This reaction is normally studied at standard temperature

and pressure. But as soon as this reaction is conducted at reduced pressure, there will

be formation of water vapour in conjunction with the hydrogen gas. Also, because the

reaction is exothermic, the released heat will elevate the kinetic energy of these products

and when passed through a diverging-converging nozzle, generate thrust.

We then looked at thermodynamic and gas laws that govern the flow of gas through a

chocked nozzle while treating the gas as ideal. This was done through specific engineering

assumptions, and the result was an ideal performance model of the propulsion system.

This led to the design of a proof of concept model of the thruster to verify the performance.

The proof of concept design provided a means of verifying the performance of the

thruster compared to the idealised model. Once the prototype thruster was designed,

an oxidiser feed system was constructed to deliver the required amount of oxidiser into

the reaction chamber. The prototype thruster was then placed on the moving plate of

thrust balance and stationed within the 𝑃𝑒𝑔𝑎𝑠𝑢𝑠 vacuum facility for in vacuum tests,

which led to a complete experimental set up. Series of experiments were conducted to

characterise the thruster. Though the experimental results were not smooth and linear

in order to predict the performance of the model, repeated experiments shows similar

trend as shown in Figure 5.1. The uncertainty in the repeated thrust measurements is

obtained as ±0.03𝑁 , while that of temperature readings is calculated as ±0.02∘C. The

uncertainty in the repeated pressure reading is obtained as ±0.6 𝑏𝑎𝑟. The uncertainty in

the pressure readings is more which is mostly due to different background pressure of

the vacuum chamber during different experiments. Other sources of errors in the course

of the experiments include:

∙ accuracy of the thrust stand

∙ accuracy of the thermocouples and pressure transducer

97

5.1. Novelty and Research Achievements

0 50 100 150 200 250 300 350 400

Time (s)

-1

0

1

2

3

4

5

6

Th

rust

(N)

# 10-3

(a) Uncertainty in thrust measurements

0 50 100 150 200 250 300 350 400

Time (s)

20

30

40

50

60

70

80

Te

mp

. (°

C)

(b) Uncertainty in temperature reading

0 50 100 150 200 250 300 350 400

Time (s)

-1.5

-1

-0.5

0

0.5

1

1.5

2

Pre

ss. (b

ar)

(c) Uncertainty in pressure readings

Figure 5.1: Uncertainty in repeated experimental data

∙ vibration in and around of the vacuum chamber and the thrust stand

∙ measuring scale

∙ unaccounted amount of oxidiser that did not get into the reaction chamber

∙ the position of the feed system from the reaction chamber

5.1 Novelty and Research Achievements

The areas of novel contributions arising from the course of this work to the field of

micropropulsion systems for CubeSats include:

∙ The novel use of the chemical combination of aluminium wool/water/sodium

hydroxide as propellant for nanosatellite propulsion.

98

5.2. Future Work

∙ Designing a low cost high impulsive CubeSat propulsion system that can deliver

high ΔV to de-orbit nanosatellite after its end of life.

5.2 Future Work

The vast literature in the area of microsatellite propulsion systems offered me the

opportunity to contribute to the dynamic field of satellite propulsion technology. However,

the research has also opened some areas that others can expend for direct continuation

of this work. These areas include and not limited to:

∙ Corrosion control: Water splitting reaction with aluminium wool using alkali salt

of sodium is safe on the containing materials if the material is non-metallic due to

corrosion. Sodium stannate salt was used in literature to reduce the associated

corrosion on the materials. In the same way, all our demonstrations in this work were

done on a reaction chamber made of stainless steel, leaving the long term aspects

of corrosion unchecked. An ideal stainless steel replacement for the qualification of

the design to higher technological level in order to mitigate against the corrosion is

polyether-ether-ketone (PEEK). It is a semi-crystalline thermoplastic material with

high performance applications in aerospace, automotive, chemical process industries

[159]. Its unique properties include outstanding resistance to chemical, solvent, fuel

and wear; insulating properties; very high temperature performance; melting point

of about 343∘C; and a high strength/weight ratio [159, 160]. The replacement of

the steel with PEEK will definitely increase the thruster performance by reducing

the dry mass ratio of the microsatellite.

∙ Addition of alcohol to the oxidizer: Water freezes in vacuum at reduced temperature

but we have performed all the experiments in a vacuum chamber whose surrounding

temperature is about 25∘C. Future work in qualifying the technology should examine

the use of alcohol (ethanol) to avoid the freezing of the oxidiser solution at low

temperature. The work should also verify the effect of alcohol to the oxidiser.

∙ Modification of the oxidizer structure: All of the experiments in this thesis relied

on a common aluminium wool with fixed fibre structure. Different structures of

the fuel (like a honeycomb structure) should be used in subsequent work, which

will provide more surface area of the fuel for reaction, and to possibly obtain a

99

5.2. Future Work

uniform level pattern of the erosion versus time.

∙ Heat losses: We had assumed that there was no heat loss to the surroundings in

analysing the heat conversion efficiency of the propulsion model. However there

could be losses through the metallic wall of the reaction chamber. These losses

should be adequately measured and accounted for in any subsequent experiments

to accurately define the efficiency of the thruster.

∙ X-ray power diffraction, 𝑋𝑅𝐷, inspection of the reaction products: Only physical

inspection of the reaction products were done at the end of every experiment for any

physical observation. A detailed understanding of the reaction and the products

could be done using X-ray power diffraction analysis.

∙ Aluminium wool: We have used an aluminium wool that is described as fine grade

and contained minimal trace of fibres [161], see the data sheet in 𝐴𝑝𝑝𝑒𝑛𝑑𝑖𝑥 B,

making it impure aluminium. Future work should try pure aluminium wool as

the propulsion fuel and compare their performances. There are also other choices

besides the nano-particles, such as honeycomb structure, pellets, rods, foils, etc.

These potential structures of aluminium should be considered in future.

100

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aluminum powder,” Materials Letters, Vol. 57, No. 19, 2003, pp. 2907–2913.

[158] VACCO, “MEPSI micro propulsion system,” http://www.cubesat-propulsion.

com/wp-content/uploads/2015/10/Mepsi-micro-propulsion-system.pdf/,

Accessed: October 2016.

[159] DirectPlastic, “All About PEEK Engineering Plastic,” http://www.

directplastics.co.uk/all-about-peek-engineering-plastic.html/, Ac-

cessed: 3, November 2015.

[160] Sakamoto, W., “Dielectric spectroscopy and thermally stimulated discharge current

in PEEK film,” Ecletica Quımica, Vol. 28, No. 2, 2003, pp. 49–53.

[161] Lustersheen-Online, “Lustersheen 1 lb. Aluminium Wool Rolls,” http:

//www.lustersheen-online.com/xcart/product.php?productid=16163&cat=

270&page=1/, Accessed: October 2016.

116

http://uk.rs-online.com/web/p/stepper-motors/5350401/

http://uk.rs-online.com/web/p/stepper-motors/5350401/

http://webbook.nist.gov/cgi/fluid.cgi?P=1&TLow=10&THigh=100&TInc=1&Applet=on&Digits=5&ID=C7732185&Action=Load&Type=IsoBar&TUnit=C&PUnit=atm&DUnit=mol%2Fl&HUnit=kJ%2Fmol&WUnit=m%2Fs&VisUnit=uPa*s&STUnit=N%2Fm&RefState=DEF/




http://www.cubesat-propulsion.com/wp-content/uploads/2015/10/Mepsi-micro-propulsion-system.pdf/

http://www.cubesat-propulsion.com/wp-content/uploads/2015/10/Mepsi-micro-propulsion-system.pdf/

http://www.directplastics.co.uk/all-about-peek-engineering-plastic.html/

http://www.directplastics.co.uk/all-about-peek-engineering-plastic.html/

http://www.lustersheen-online.com/xcart/product.php?productid=16163&cat=270&page=1/



Appendix A

Detail Drawings of the Hybrid

Propulsion Thruster

117

APPENDIX A. DETAIL DRAWINGS OF THE HYBRID PROPULSION THRUSTER

5lluLenghtsuinumm5nglesu±7°uDistanceu±u7

SIZE@A4PartuName@uExpCombustionChberNowBdft

Scale@u7u@u7

Material@

PartsuRequired@

SurreyuSpaceuCentreuySSC(UniversityuofuSurreyvu95u9uilding

GuildfordvuSurreyvuGUXu4XH

X2G26GX2705hmeduOzomatauDavid

oa227743surreyBacBuk

2402076O726

SteelXu7

ProjectuSheet@

5DET5ILu5

u

99

SECTIONu9,9u

C

DET5ILuC

O0H

O02O

HX

OH2

NonuThreadeduforuMXuXu6

62°

X6

6vH/

0H

zz

zzv27

7

H

72vzOX/v47

7vH

HX6vH/

6vH/

O6vH/

Figure A.1: Reaction chamber

118


All.Lenghts.in.mmAngles.±7°.Distance.±.7

SIZEkA4Part.Namek.Exp_NozzlePartBdft

Scalek.7.k.7

Materialk

Parts.Requiredk

Surrey.Space.Centre.CSSC(University.of.Surreyv.BA.Building

Guildfordv.Surreyv.GUX.4XH

X2G26GX270Ahmed.Ozomata.David

oa227749surreyBacBuk

2402076O726

SteelX.7

7/

0/°O

62

X647°

Non.Threaded.for.MX.Bolts.and.Nuts.X.6

Project.Sheetk

OH2

62H2

7

@vX/

X2 77v/

Figure A.2: Nozzle part 1

119


AllfLenghtsfinfmmAnglesf±X°fDistancef±fX

SIZE6A4PartfName6fNozzleNutGdft

Scale6f/f6fX

Material6

PartsfRequired6

SurreyfSpacefCentref(SSC)UniversityfoffSurrey,fBAfBuilding

Guildford,fSurrey,fGUHfOXH

7@27O2H7X8AhmedfOzomatafDavid

oa77XXO@surreyGacGuk

7O878X4@X74

SteelXfX

ProjectfSheet6

7,O

XX

(From

Sw

agelok)

Figure A.3: Nozzle part 2 (Modified from swagelok [143])

120


AlldLenghtsdindmmAnglesd±H°dDistanced±dH

SIZE8A4PartdName8dExpPlatformHGdft

Scale8dHd8d4

Material8

PartsdRequired8

SurreydSpacedCentred(SSC)UniversitydofdSurrey,dBAdBuilding

Guildford,dSurrey,dGU4d@XH

H07X@74XH0AhmeddOzomatadDavid

oaXXHH@@surreyGacGuk

X@0X0HzkHXz

SteelXdH

ProjectdSheet8

A

DETAILdAd

0X

/X/X/XHO

46

HXk,O

46MzdThreaded

HH,O

H06,O

/4z

ThreadeddfordM0dBoltsdXdz

O0X

HoledfordMzdBoltsdanddNuts

Figure A.4: Thrust attachment to the thrust stand

121

Appendix B

Experiment Hardware

122

APPENDIX B. EXPERIMENT HARDWARE

CALL US: 1 - 8 0 0 - 2 7 4 - 9 2 9 9

LusterSheen-Online : : Metallic Wools : : Alum inum Wool : : Lustersheen 1 lb. Alum inum Wool Ro lls

Lu st er sh een 1 lb . A lu m in u m W oo l Ro l l s

Alum inum bulk ro lls are 4" w ide w ith the thickness being approxim ately 1/ 4" in anuncom pressed state....length varies at about 18 ' feet long w hen unrolled.

Made in the USA

Lustersheen 1 lb ro lls are m ade of Alum inum Alloy: AA 5056 and are a USA product .

Det a i ls

SKU: SKU161633Shipping Weight : 1.38 lbs

Pr i ce : $ 2 4 .0 0

Op t ion s

Grades

Quant it y

Add t o car t

Recom m en d ed p r od u ct s l i st

Custom ers who bought this product also bought the fo llow ing products:

A I SI 4 3 4 St ain less St eel Wool 1 lb . Ree lsLust ersheen A lum inum Wool RibbonsLust ersheen Copper WoolXcluder 1 " x 4 ' foo t St r ipsClassic Ter racot t a Wax , 1 lb sizeAI SI 3 1 6 L St a in less St ee l Wool 1 lb ro llLust e rsheen St r ipping Mesh3 Pack Bronze Wool PadsLust ersheen Brass Woo lXcluder St ar t er K it

Powered by X-Cart ecom m erce softw are Copyright © 2008-2016 LusterSheen-Online

LusterSheen-Online :: Metallic Wools :: Aluminum Wool :: Lustershee... file:///C:/Users/Public/Documents/LusterSheen-Online Metallic Wools...

1 of 1 20/10/2016 13:07

Figure B.1: Aluminium wool data sheet

123


RS Part No’s: See Below

Page 1 of 3

Data sheetMineral Insulated ThermocouplesTypes ‘K’ or ‘J’ with 1 metre lead & tails – stainless steel sheath

RS,jAx°b,kjF

8Note+ Illustration shows Type ‘K’f

Mineral insulated Thermocouple to IEC A&F

Choice of Type ‘K’ with °j, stainless steel sheath or Type ‘J’ with °Pj stainless steel sheath

Highly flexibleH sheath can be bentbformed to suit many applications and processes

Insulated hot junction

Plain pot seal 8P,,°Cf

j metre kb,xPmm PFA Teflon® insulated flat pair cable and tails 8colour coded to IEC A&Ff

Specifications

Sensor type+ Type ‘K’ 8Nickel ChromiumbNickel AluminiumfType ‘J’ 8IronbConstantanf

Construction+ Flexible mineral insulated probe with stainless steel sheathH plain pot sealE j metre extension cable

Elementbhot junction+ Single elementH junction insulated from sheath 8offers protection againstspurious electrical signalsf

Termination+ j metre kb,xPmm PFA Teflon® insulated flat pair cableH colour coded inaccordance with IEC A&F

Probe temperature range+ Type ‘K’ gF,°C to 2jj,,°C >jx,mm diametergF,°C to 2kA,°C – jx,mm diameter and below

Type ‘J’ gF,°C to 2kA,°CPot seal rating+ P,,°C

Figure B.2: K-type insulated thermocouple data sheet

124


www.swagelok.com

4RP Series High-purity PFA material

Working pressures up to 180 psig (12.4 bar)

Temperatures up to 300°F (148°C)

1/4, 3/8, and 1/2 in. Swagelok® PFA tube fitting end connections

PFA Need le Valves

Figure B.3: PFA needle valve data page 1

125


2 PFA Needle Valves

Features Straight-through orif ce for full f ow

Plug stem tip for flow regulation

Low-torque, leak-resistant stem seal

High-purity PFA material with low extractables for chemical resistance

Swagelok PFA tube fitting end connections for consistent performance

Technical Data

Pressure-Temperature Ratings

TestingEvery 4RP series PFA needle valve is factory tested with nitrogen at its rated pressure for leakage at the seat to a maximum allowable leak rate of 0.1 std cm3/min. The stem seal is tested with helium at rated pressure to a maximum leak rate of 1 3 10–3 std cm3/s.

Cleaning and PackagingEvery 4RP series PFA needle valve is cleaned and packaged inaccordance with Swagelok Standard Cleaning and Packaging (SC-10), MS-06-62.

Oval handle for easy actuation

Bonnet with positive stem stop prevents accidental disassemblyAcme stem threads

add strength

Plug stem tip ensures leak-tight shutoff

Straight-through orifice permits full flow

Panel mounting

Stem seal design provides reduced operating torque

Swagelok Tube Fitting End Connection

Ordering Number

Orifice in. (mm) Cv

1/4 in. PFA-4RPS4 0.156 (4.0) 0.38

3/8 in. PFA-4RPS6 0.250 (6.4) 1.39

1/2 in. PFA-4RPS8

End Connection 1/4 in., 3/8 in. 1/2 in.

Temperature, °F (°C) Working Pressure, psig (bar)

0 (–17) to 70 (21)100 (37)150 (65)

180 (12.4)160 (11.0)125 (8.6)

125 (8.6)110 (7.5)

87 (5.9)

200 (93)250 (121)300 (148)

95 (6.5)69 (4.7)50 (3.4)

66 (4.5)49 (3.3)33 (2.2)

A packing adjustment may be required periodically to increase service life and to prevent leakage.

Valves that have not been cycled for a period of t ime may have a higher init ial actuat ion torque.

To increase service life, ensure proper valve performance, and prevent leakage, apply only as much torque as is required to achieve posit ive shutof f.


126


PFA Needle Valves 3

Materials of Construct ion

Wetted components listed in italics.

➀ Blue dyed.

Flow Data at 70°F (20°C)

PFA-4RPS6 and PFA-4RPS8

PFA-4RPS4

Number of Turns Open

Flow

Coe

ffici

ent (

Cv)

Number of Turns Open

Flow

Coe

ffici

ent (

Cv)

Flow Coefficient at Turns Open

Ordering Number

Pressure Drop to Atmosphere ( p)

psi (bar)Air Flow

std ft3/min (std L/min)Water Flow

U.S. gal/min (L/min)

PFA-4RPS4

10 (0.68) 4.3 (120) 1.2 (4.5)

75 (5.1) 15 (420) 3.3 (12)

180 (12.4) 34 (960) 5.1 (19)

PFA-4RPS6

10 (0.68) 15 (420) 4.4 (16)

75 (5.1) 57 (1600) 12 (45)

180 (12.4) 120 (3300) 18 (68)

PFA-4RPS8

10 (0.68) 15 (420) 4.4 (16)

75 (5.1) 57 (1600) 12 (45)

125 (8.6) 90 (2500) 15 (56)

Component Material Grade/

ASTM Specification

1 Handle insert➀

PFA 440-HP/ D3307

2 Screw➀

3 Handle➀

4 Packing nut➀

5 Upper packing

6 Lower packing

7 Stem stop

8 Panel nut➀

9 Stem

10 Body, ferrules

11 Nut➀

Lubricant PTFE based

1

2

3

4

5

6

7

8

9

10

11


127


Safe Product SelectionWhen selecting a product, the total system design must be considered to ensure safe, trouble-free performance. Function, material compatibility, adequate ratings, proper installation, operation, and maintenance are the responsibilit ies of the system designer and user.

Caut ion: Do not mix or interchange parts with those of other manufacturers.

Warranty InformationSwagelok products are backed by The Swagelok Limited Lifetime Warranty. For a copy, visit swagelok.com or contact your authorized Swagelok representative.

Ordering Information and DimensionsSelect an ordering number.

Dimensions, in inches (millimeters), are for reference only and are subject to change.

Panel Hole Drill Dimensions

Maximum panel thickness is 0.25 in. (6.4 mm).

3.34 (84.8) open

1.95 (49.5)

2.00 (50.8)

0.625 (15.9)

ADimensions shown with Swagelok nuts positioned prior to swaging.

0.19 (4.8)

1.19 (30.2)

1.00 (25.4)

0.88 (22.2)

or0.88 (22.2)

Installat ion PFA tubing MUST be grooved for use with PFA tube f t t ings. Use the Swagelok groove cut ter tool.

3. Continue tightening until the nut and body hexes are aligned.

2. While holding f tting body steady, tighten the blue nut until there is no gap between the nut and body hexes.

1. Insert grooved PFA tubing into the Swagelok PFA tube f tting until a clicking sound is heard.

Ordering Number

A in. (mm)

Weight lb (kg)

PFA-4RPS4 2.50 (63.5) 0.20 (0.09)

PFA-4RPS6 2.66 (67.6) 0.23 (0.10)

PFA-4RPS8 3.14 (79.8) 0.27 (0.12)

Swagelok—TM Swagelok Company© 2001–2013 Swagelok CompanyPrinted in U.S.A., AGSMarch 2013, R8MS-01-69

0.19 (4.8)

1.19 (30.2)

Other Swagelok PFA ProductsFor more information about Swagelok PFA tubing and tools, see the Swagelok Hose and Flexible Tubing catalog, MS-01-180.

For more information about Swagelok PFA tube f ttings, see the Swagelok PFA Tube Fittings catalog, MS-01-05.

For more information about Swagelok PFA plug valves, see the Swagelok PFA Plug Valves catalog, MS-01-56.

PFA Plug ValvePFA Tubing PFA Tube Fit t ings


128


Data SheetRS stock number 535-0401

Data Pack Issued July 2006 1504256272

42mm 1.8' HIGH TORQUE STEPPER200 STEP

Specificat ionsModel 535-0401

Step Angle 1.8°

Step Angle Accuracy(Full Step, No Load)(%)

±5%

Rated Voltage (V) 2.8

Current/ Phase (A) 1.68

Resistance/Phase (Ω) 1.65

Inductance/Phase (mH) 2.8

Detent Torque (mNm) 25

Holding Torque (Ncm) 44

Rotor Inertia (g-cm²) 68

Weight (Kg) 0.35

Number Of Leads (No.) 4

Characterist icsResistance Accuracy

±10%

Inductance Accuracy±20%

Temperature Rise80ºC Max. (Rated Current 2 phase On)

Ambient Temp-20 ºC to +50 ºC

Insulation Resistance100mm Ωmin.,500Vdc

Dielectric Strength500Vac for 1 min

Shaft Radial Play0.06mm Max. (450G-Load)

Shaft Axial Play0.08mm Max. (450G-load)

RSOComponentsOshallOnotObeOliableOforOanyOliabilityOorOlossOofOanyOnatureO:howsoeverOcausedOandOwhetherOorOnotOdueOtoORSOComponents’Onegligence0OwhichOmayOresultOfromOtheOuseOofOanyOinformationOprovidedOinORSOtechnicalOliterature4

RSOComponents6OPOOBoxO996OCorby6ONorthants6ONN©7O9RSO Telephone:OE©536O8E©834

AnOElectrocomponentsOCompanyO ©ORSOComponentsO©998

Dimensions in mm

Speed V Torque Characterist ics

Figure B.7: DC stepper motor data sheet

129


CapsbandbPlugs

PartbNox

PartbDescription:

M537bCapbforbA94binxbODbTubing

M537q433qC

Body Material Super Duplex Stainless Steel

Cleaning Process Standard Cleaning and Packaging (SC-10)

Configuration Cap

Connection 1 Size 1/4 in.

Connection 1 Type Swagelok® Tube Fitting

eClass (4.1) 37020713

eClass (5.1.4) 37020517

eClass (6.0) 22-56-02-07

eClass (6.1) 37-02-05-17

UNSPSC (10.0) 40142607

UNSPSC (11.0501) 40142607

UNSPSC (13.0601) 40183104

UNSPSC (15.1) 40183104

UNSPSC (4.03) 40141706

UNSPSC (SWG01) 40141706

ProductbSpecifications

©bM3A5bSwagelokbCompany

ThebcompletebcatalogbcontentsbmustbbebreviewedbtobensurebthatbthebsystembdesignerbandbuserbmakebabsafebproductbselectionxbWhenbselectingbproductsFbthebtotalbsystembdesignbmustbbebconsideredbtobensurebsafeFbtroubleqfreebperformancexbFunctionFbmaterialbcompatibilityFbadequatebratingsFbproperbinstallationFboperationFbandbmaintenancebarebthebresponsibilitiesbofbthebsystembdesignerbandbuserx

Caution:bDobnotbmixborbinterchangebvalvebcomponentsbwithbthosebofbotherbmanufacturersx

General

A39A49M3A5b9:33:37bAM

swagelokxcom

Figure B.8: Swagelok cap and plug data sheet

130


GEM-SOL Chem-Sol 1/4"2/2 Way NC , NO

How to OrderExample : GEM-C-1201V1-321Is a GEM-SOL Chem-Sol, 1/4"BSP, 2W NC with Viton,plastic manual override, 24V AC 8W 60Hz with connector.

General Description

These GEM-SOL 2/2 way NC, NO Chem-Solsolenoid valves are recommended for use inapplications where corrosive fluid must becontrolled, such as chemical process, watertreatment, analysis device etc.

They can be used for industrial and irrigationcontrol and automation systems.

Notes

Contact our technical department to getdetails on valve and fluid compatibility.To order valves manufactured to yourspecific requirements, please contact ourtechnical department.ADC valves are suitable to work only withAC 8W or DC 10W coils.

BACCARA

1.55

FunctionBodyGEM-C Seals ManualOverridePort

2021

1/4"BSP1/4"NPT

PPA 1 VITONEPDMSilicone

12

2W NC2W NO

NonePlastic

01

VES

Voltage ConnectorPower

W/out coil6122448110120230240other

0123456789

No coilAC8W 50HzAC8W 60HzDC10WAC5.5W 50HzAC5.5W 60Hz

DC 5.5W

withoutwithwith LEDwith bi colorLEDflying leadscoilwith 1/2" Hub

0123457

0123

4

5

NCNO

Flow direction

Figure B.9: Solenoid valve description and data sheet 1

131


Technical Specifications

Dimensions

1.56 SV-3-05

Functions: 2/2 Way NC or NO

Ports size: 1/4" BSP & NPT

Orifice: 4.5mm

Kv: 5 L/min

Pressure range: See table

Temperature range: Fluid: -15 C to max 90 C

Ambient: -10 C to 50 C

Manual Override: NC: Plastic. The coil can be rotated in

4 positions, each 90

NO: Without manual override

Materials in contact

with fluid: Main valve :

Reinforced PPA

Diaphragm:

Viton, EPDM, Silicone

Weight (with coil): 200 gr

Coil voltage: All Baccara coils voltages 10%

AC 8W, 5.5W

DC 10W, 5.5W

Protection class IP65 with plug attached

O

O

O

Pressure Table - NC

CoilFlow directionnot restricted

ADC

AC 8W

DC 10W

AC 5.5W

DC 5.5W

+-

O

O

Flow directionrestricted

0-2 bar

0-2.5 bar

0-2.5 bar

0-2 bar

0-1.5 bar

0-0.5bar

0-0.7 bar

0-0.7 bar

0-0.5 bar

0-0.4 bar

Pressure Table - NOCoil and

power ratingFlow directionnot restricted

ADCAC 8w orDC 10W

Flow directionrestricted (1)

0 - 1 bar 0 - 0.5 bar

5.5WAC/DC 0 - 1 bar 0 - 0.5 bar

(1) Higher input presssure of up to 1atm can be achieved with minimal pressure drop on the valve of 0.3 atm.

Figure B.10: Solenoid valve description and data sheet 2

132


ModelLcNzzzz;j•ShippingcWeightLcz,zj;Kg•

N

Arduino Uno SMD Rev3

ThecNrduinocUnocSMDcRjciscacmicrocontrollercboardcbasedconcthecNTmegajB-c4datasheet/,cItchascykcdigitalcinputHoutputcpinsc4ofcwhichcvccancbecusedcascPWMcoutputs/6cvcanalogcinputs6cacyvcMHzccrystalcoscillator6cacUSBcconnection6cacpowercjack6cancICSPcheader6candcacresetcbutton,cItccontainsceverythingcneededctocsupportcthecmicrocontrollerEcsimplycconnectcitctocaccomputercwithcacUSBccablecorcpowercitcwithcacNCPtoPDCcadaptercorcbatteryctocgetcstarted,

ThecUnocdifferscfromcallcprecedingcboardscincthatcitcdoescnotcusecthecFTDIcUSBPtoPserialcdrivercchip,

NdditionalcfeaturesccomingcwithcthecRjcversioncareL

NTmegayvUBcinsteadc-UBcascUSBPtoPSerialcconverter,•y,zcpinoutLcaddedcSDNcandcSCLcpinscforcTWIccommunicationcplacedcnearctocthecNREFcpincandctwocothercnewcpinscplacedcnearctocthecRESETcpin6cthecIOREFcthatcallowcthecshieldsctocadaptctocthecvoltagecprovidedcfromcthecboardcandcthecsecondconeciscacnotcconnectedcpin6cthatciscreservedcforcfuturecpurposes,

•

strongercRESETccircuit,•c2Uno2cmeansc2One2cincItaliancandciscnamedctocmarkcthecupcomingcreleasecofcNrduinocy,z,cThecUnocandcversioncy,zcwillcbecthecreferencecversionscofcNrduino6cmovingcforward,cThecUnocisctheclatestcincacseriescofcUSBcNrduinocboards6candcthecreferencecmodelcforcthecNrduinocplatformEcforcaccomparisoncwithcpreviouscversions6cseectheindexcofcNrduinocboards,cTechnical Specifications

Microcontroller NTmegajB-OperatingcVoltage CVSupplycVoltagec4recommended/ ;PyBVMaximumcsupplycvoltagec4notcrecommended/BzVDigitalcIHOcPins ykc4ofcwhichcvcprovidecPWMcoutput/NnalogcInputcPins vDCcCurrentcpercIHOcPin kzcmNDCcCurrentcforcj,jVcPin CzcmNFlashcMemory jBcKBc4NTmegajB-/cofcwhichcz,CcKBcusedcbycbootloaderSRNM BcKBc4NTmegajB-/EEPROM ycKBc4NTmegajB-/

ClockcSpeedcyvcMHzc

IfcyoucwantctocgivecaccloserclookctocthiscboardcwecadvicecyouctocvisitcthecofficialcNrduinocUNOcpagecincthecHardwarecSection,

Figure B.11: Arduino Uno SMD Rev3 data sheet

133


16

3x Mounting holesø4.5 mm 3x Mounting

holesø4.5 mm

MR

optoNCDT 1700 (2/10/20/50/100/200/250VTmm)

SMR

MR

optoNCDT 1700 (40/500/750mm)

Start of measuringrange

End of measuringrange

End of measuringrange

SMR

97 150

75

140130

A

A

B

B

12

13.4

24.2

31 75

17.5

36.1

17.5

37.5

13.2

15

30

35

67

5

75

18.5

40

70 80

4

8980

ø4

ø5

ø8

15

15

α

ϕ

εα

ϕ

εStart of measuringrange

(Dimensions in mm, not to scale. All CAD files are available online.)

MR SMR α ϕ ε A B

2 24 35° 40° 44.8° 25.8 16.8

10 30 34.3° 35.2° 35.6° 28.7 20.5

20 40 28.8° 27.5° 26.7° 30.1 22.0

50 45 26.5° 23.0° 18.3° 31.5 22.5

100 70 19.0° 15.4° 10.9° 32.6 24.1

200 70 19.0° 9.78° 6.97° 33.1 24.1

250VT 70 19.0° 8.4° 6.0° 33.5 24.1

40 175 22.1° 21.9° 21.8° 101 86

500 200 19.3° 9.8° 7.0° 101 85

750 200 19.3° 7.7° 5.0° 101 85

The benchmark in laser triangulation sensorsThe optoNCDT 1700 series is truly a world leading laser displacement sensor. Featuring Real Time Surface Compensation (RTSC), remote software programming and excellent linearity & resolution the optoNCDT 1700 is difficult to match at this price level. Integrated conditioning electronics allows the sensor to have a very unique and compact design.

Adjustable exposure time/measuring rateFor poor reflecting targets, the measuring rate can be reduced to enable a longer exposu-re time. The set measurement rate always remains constant so that with closed-loop control the system response time is always the same.

Adjustable limit switchesAs well as for precise measurement, the optoNCDT 1700 sensors are also used for tolerance or limit monitoring. Two switching points are available which can be configured and adjusted via the remote software (USB connection). The switching hysteresis can also be individually adjusted for each limit point.

AnalogDigital

Analogue (U/I)and digital output

Calibration certificate included

Adjustable filter functions(firmware)

Filter inside

High flex cables for dragchain or robot use

Adjustable measuring rateup to 2.5kHz

312Hz375Hz

1000Hz

Real Time Surface Compensation

Eleven models with measuringranges from 2mm to 1000mm

14-pin-connector(Pin side female cable connector or solder-pin side male cable connector)

Connector (sensor side)Article Number: 0323272

ø15

~ 50

Connector (sensor cable)Article Number: 0323243

~ 51

~15

Sensor with integrated controller for industrial applications optoNCDT 1700

Figure B.12: Laser displacement sensor (optoNCDT 1700-50) data page 1

134


17

ModelILD

1700-2ILD

1700-10ILD

1700-20ILD

1700-40ILD

1700-50ILD

1700-100ILD

1700-200ILD

1700-250VTILD

1700-500ILD

1700-750

Measuring range 2mm 10mm 20mm 40mm 50mm 100mm 200mm 250mm 500mm 750mm

Start of measuring range 24mm 30mm 40mm 175mm 45mm 70mm 70mm 70mm 200mm 200mm

Midrange 25mm 35mm 50mm 195mm 70mm 120mm 170mm 195mm 450mm 575mm

End of measuring range 26mm 40mm 60mm 215mm 95mm 170mm 270mm 320mm 700mm 950mm

Linearity2µm 8µm 16µm 32µm 40µm 80µm 200µm 630µm 400µm 750µm

FSO ≤ 0.1% ≤ 0.08% ≤ 0.1% ≤ 0.25% ≤ 0.08% ≤ 0.1%

Resolution (at 2.5kHz without averaging)

0.1µm 0.5µm 1.5µm 4µm 3µm 6µm 12µm 50µm 30µm 50µm

Measuring rate 2.5kHz / 1.25kHz / 625Hz / 312.5Hz (adjustable)

Light source semiconductor laser < 1mW, 670nm (red)

Permissable ambient light (at 2.5kHz) 10,000lx 15,000lx 10,000lx

Laser safety class class 2 acc. DIN EN 60825-1 : 2008-05

Spot diameter

SMR 80µm 110µm 320µm 230µm 570µm 740µm 1300µm 1500µm 1500µm 1500µm

MMR 35µm 50µm 45µm 210µm 55µm 60µm 1300µm 1500µm 1500µm 1500µm

EMR 80µm 110µm 320µm 230µm 570µm 700µm 1300µm 1500µm 1500µm 1500µm

Temperature stability1) 0.025%FSO/°C

0.01 % FSO/°C0.025%FSO/°C

0.01 %FSO/°C

Operation temperature 0 ...+ 50°C 0 ...+ 55°C 0 ...+ 50°C

Storage temperature -20 ... + 70°C

Outputmeasurements selectable: 4 ... 20mA / 0 ... 10V / RS 422 / USB (optional with cable PC1700-3/USB)

switching outputs 1 x error or 2 x limit (each pogrammable)

Switch Input laser ON-OFF / zero

Operation via touch screen on sensor or via PC with ILD 1700 tool

Power supply 24VDC (11 ... 30VDC), max. 150mA

Electromagnetic compatibility (EMC) EN 61000-6-3 EN 61000-6-2

Sensor cable length (with connector) 0.25m (integrated cable with connector) option: 3m or 10m

Synchronisation possible for simultaneous or alternating measurements

Protection class IP 65

Vibration 2g / 20 ... 500Hz

Shock 15g / 6ms

Weight (with 0.25m cable) ~ 550g ~ 600g ~ 550g ~ 600g

FSO = Full Scale Output All specifications apply for a diffusely reflecting white ceramic target1) based on digital outputSMR = Start of measuring range MMR = Midrange EMR = End of measuring range

Custom Sensor ModificationsFor applications where the above standard sensors do not meet your requirements, it may be possible to supply asensor with modified specification. Please contact us for further information.

Options Non standard measuring range and stand off Custom housing or mounting geometry Non standard signal interfaces Special cable length of electrical connector 90° beam deflection

Vacuum suitability Reduced mass Increased shock and vibration resistance

Figure B.13: Laser displacement sensor (optoNCDT 1700-50) data page 2

135

Appendix C

Program Codes

C.1 Solenoid Valves Control Program

/*

So leno id va lve s c on t r o l program to f i l l plenum volume with

o x i d i s e r and empty i t i n to the r e a c t i on chamber that

k i ck s t a r t the chemica l r e a c t i on in the r e a c t i on chamber

*/

i n t ValvePin1 = 4 ; // So l eno id va lve1

connected to pin 4

i n t ValvePin2 = 5 ; // So l eno id va lve2

connected to pin 4

i n t SwPin = 12 ;

i n t buttonWas = 0 ; // The s t a t e o f the switch ( pushed = 1 ,

not pushed = 0) l a s t time we looked

i n t buttonIs = 0 ; // Current s t a t e o f the switch

i n t LEDPin1 = 13 ;

i n t LEDPin2 = 11 ;

void setup ( ) // run once , when the sketch s t a r t s

pinMode ( ValvePin1 , OUTPUT) ; // s e t s the d i g i t a l pin

as output

136

C.1. Solenoid Valves Control Program

pinMode ( ValvePin2 , OUTPUT) ; // s e t s the d i g i t a l pin

as output

pinMode (SwPin , INPUT) ;

pinMode (LEDPin1 , OUTPUT) ;

pinMode (LEDPin2 , OUTPUT) ;

buttonIs = d ig i t a lRead (SwPin ) ; //Read the i n i t i a l s t a t e

o f the switch !

//===========Functions=====================//

void getButton ( )

buttonWas = buttonIs ; // Set the o ld s t a t e o f the button

to be the cur rent s t a t e s i n c e we ’ re c r e a t i n g a

new current s t a t e .

buttonIs = d ig i t a lRead (SwPin ) ; // Read the button s t a t e

void openValve1 ( )

d i g i t a lWr i t e ( ValvePin1 , HIGH) ;

d i g i t a lWr i t e (LEDPin1 ,HIGH) ;

de lay (4000) ; // Wait f o r the plenum volume to f i l l up

d i g i t a lWr i t e ( ValvePin1 , LOW) ;

d i g i t a lWr i t e (LEDPin1 ,LOW) ;

// de lay (2000) ; // Wait f o r va lve1 to s e t t l e down be f o r e

va lve2 opens

void openValve2 ( )

d i g i t a lWr i t e ( ValvePin2 , HIGH) ;

d i g i t a lWr i t e (LEDPin2 ,HIGH) ;

de lay (10000) ; // Wait f o r the plenum volume to empty



void c l o s eVa lve ( )


137

C.2. Thrust Balance Calibration Constant Program




//=================Main Loop===================//

void loop ( ) // run over and over again

getButton ( ) ;

i f ( ( buttonIs==1)&&(buttonWas==0))

openValve1 ( ) ;

i f ( ( buttonIs==1)&&(buttonWas==0))

openValve2 ( ) ;

e l s e

c l o s eVa lve ( ) ;

C.2 Thrust Balance Calibration Constant Program

% The program was adopted from Dr Char l i e Ryan f o r t h i s work

c l e a r a l l

c l o s e a l l

% path to data , ending in backs la sh

path=’C:∖ Users ∖oa00117∖Documents∖START TO WRITE∖HRM Experiments

∖ThrustBal Spring 2015∖ Ca l i b ra t i on Constant F i l e s ∖ ’ ;

% [ FileName , PathName ] = u i g e t f i l e ( ’ * . csv ’ , ’ S e l e c t the MATLAB

code f i l e ’ ) ;

% sample ra t e w i l l normally be 312 .5 , or 2500Hz

samplerate =2500;

138


l =265; %length o f the s t r i n g on which the weight was suspended

( a l l ow ing f o r

%d i s t anc e tothe cent r e o f the weight .

d=15; %l a t e r a l d i s t a c e over which the wieght was d i sp l a c ed (mm)

m=19.49E−3; %mass o f the weight used (Kg)

g=9.81; %g r a v i t a t i o n a l a c c l e r e a t i o n at sea l e v e l

f=(d/ l ) *m*g ; % the f o r c e which was app l i ed to the thrus t

ba lance

% a vec to r i s f i l e s to skip− sk ipped f i l e s w i l l not appear in

the average

skipped = [ ] ;

f i g u r e (99)

%which f i l e number to open − ’ mu l t i s t e p s c a l i b r a t i o n n ’

n=14;

%f i g u r e

%load raw data−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−

number = num2str ( n ) ;

f i l e = s t r c a t ( path , ’ mu l t i s t e p s c a l i b r a t i o n ’ , number , ’ .

x l sx ’ ) ;

RawData = load ( f i l e ) ;

f i l e ;

%−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−

x = l i n s p a c e (1 , l ength (RawData) . / samplerate , l ength (RawData) ) ;

% 1 to go 0 to not go

i f 0

newStartTime = 100 ;

indx1 = f i nd (x>newStartTime , 1 ) ;

139


x = x( indx1 : end ) ;

RawData = RawData( indx1 : end , : ) ;

end

%apply butterworth f i l t e r to cur rent data s e t

[ b , a ] = butte r ( 3 , 0 . 0 003 , ’ low ’ ) ;

s i g n a l l ow = f i l t f i l t (b , a , RawData ( : , 2 ) ) ;

%s i gna l l ow (1) =0;

f i g u r e

%p lo t (RawData ( : , 1 ) ,RawData ( : , 2 ) )

hold a l l

% OPTIONAL: p l o t each f i l t e r e d data s e t

p l o t ( l i n s p a c e (0 , l ength ( s i g na l l ow ) / samplerate , l ength ( s i g na l l ow

) ) , s i gna l l ow , ’ k ’ , ’ LineWidth ’ , 2 )

%f i nd d e r i v a t i v e o f the data

y= d i f f ( s i gna l l ow , 1 ) ;

% chooses the thresho ld , g i v ing the po in t s at which the s i g n a l

drops / r i s e s

% from the d i f f e r e n t i a t i o n o f the s i g n a l . Like a standard

dev i a t i on ( sigma )

% value

%typ i c a l va lue i s 5 or 10

chosenMedThresh = 20 ;

%

% f ind thre sho ld

medThresh = median ( abs (y ) ) ;

% make element > thresh *7 = 1

yCross = ( abs (y )>medThresh .* chosenMedThresh ) ;

140


% get c r o s s i n g s as + or −1

[ s tart Indx , yVal ] = f i nd ( d i f f ( yCross )==1) ;

[ endIndx , yVal ] = f i nd ( d i f f ( yCross )==−1) ;

%adds an add i t i o na l c r o s s i n g po int at the s t a r t

s ta r t Indx = [ s ta r t Indx ; l ength (y ) ] ;

%adds and add i t i ona l c r o s s i n g po int at the end o f the data

endIndx = [ 1 ; endIndx ] ;

f i g u r e (97)

p l o t ( abs (y ) )

hold on

p lo t ( y.*0+medThresh .* chosenMedThresh , ’ r ’ )

g r i d

f i g u r e

p l o t (x , RawData ( : , 2 ) , ’ r ’ )

hold on

p lo t (x , s i gna l l ow , ’ b ’ )

g r id on

f o r iS tep = 1 : l ength ( s ta r t Indx )

% get va lue s between end and s t a r t

s tepVals = s i gna l l ow ( endIndx ( iS tep ) : s t a r t Indx ( iS tep ) ) ;

s t epResu l t s ( iS tep ) = median ( s tepVals ) ;

p l o t (x , s i g n a l l ow .*0+ s t epResu l t s ( iS t ep ) , ’ g ’ )

end

% average va lue s

sepDis t = d i f f ( s t epResu l t s ) ;

midPoint = f l o o r ( l ength ( sepDis t ) /2) ;

sepVec1 = sepDis t ( 1 : midPoint ) ;

sepVec2 = abs ( sepDis t ( end :−1:midPoint+1) ) ;

f o r iForce= 1 : l ength ( sepVec1 ) ;

141

C.3. Thrust Response Program of One-shot Experiment

f o r c e ( iForce )=iForce * f ;

end

f i g u r e (99)

%p l o t s out va lue s o f the f o r c e ver sus the cummulative add i t i on

o f the drop or r i s e

% sepVec1 i s the drop down

% sepVec2 are the r i s e s

p l o t ( f o r c e , cumsum( sepVec1 ) , ’ o ’ )

hold on

p lo t ( f o r c e , cumsum( sepVec2 ) , ’ s ’ )

y l ab e l ( ’ d i sp lacement (mm) ’ )

C.3 Thrust Response Program of One-shot Experiment

%This program i s used to f i nd the thrus t re sponse o f one−shot

experiment o f the propu l s i on model

c l e a r a l l

c l o s e a l l

%load raw data−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−

x = importdata ( ’ExpN. txt ’ ) ;% c a l l the exper imenta l data , N i s

the experiment number

df = 0.358792849;% Ca l i b ra t i on Constant

RawData = (−(x ( : , 2 )−x (1 , 2 ) ) .* df ) ; %making the s t a r t i n g po int

0 and i nv e r s i n g i t

%−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−

%% F i l t e r the data

fNorm = 0 .5/ ( 312 . 5/2 ) ; % 312 .5 i s the sample ra t e

[ b , a ] = butte r (3 , fNorm , ’ low ’ ) ; % vary the number 1 to change

the amount o f f i l t e r i n g app l i ed to the raw data : h igher

number=gr ea t e r f i l t e r i n g

s i g na l l ow = f i l t f i l t (b , a , RawData) ;

142

C.4. Thrust Response Program of Repeat Cycle Injection

xTime = x ( : , 1 )−x (1 , 1 ) ;% making the s t a r t i n g time 0

%In t e g r a l = trapz (max( s i gna l l ow , 0 ) ) ;

%%Remove the non−p o s i t i v e data

[ S , ˜ ] = s i z e ( s i g na l l ow ) ;

[ S , ˜ ] = s i z e (xTime) ;

f o r s = 1 : S

i f ( s i g n a l l ow ( s , : ) < 0) ;

s i g n a l l ow ( s , : ) = 0 ;

end

end

%%Plot the data

p l o t (xTime , s i gna l l ow , ’ k ’ )

ax i s ( [ 0 200 0 0 . 0 4 ] )

%gr id on ;

%t i t l e ( ’ Thrust performance ’ )

x l ab e l ( ’ Time ( s ) ’ ) ; y l ab e l ( ’ Thrust (N) ’ ) ;

%%Find the t o t a l impulse o f the re sponse

Tota l Impulse = trapz (xTime , s i g n a l l ow )

C.4 Thrust Response Program of Repeat Cycle Injection

% This program i s used to f i nd the thrus t re sponse o f repeat

c y c l e s o f the propu l s i on model

c l e a r a l l

c l o s e a l l

%load raw data−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−

x = importdata ( ’ExpN. txt ’ ) ; % c a l l the exper imenta l data , N i s

the experiment number

143


a= x ( : , 1 ) − x (1 , 1 ) ; % make the s t a r t i n g time zero

df =0.326761184; % Ca l i b ra t i on Constant

b = (−( x ( : , 2) − x (1 , 2 ) ) .* df ) ; % inv e r s e the data , make i t

s t a r t from zero and mul t ip l e i t by the c a l i b r a t i o n constant

%% F i l t e r the data

fNorm = 0 .5/ ( 312 . 5/2 ) ;

[ b1 , a1 ] = butte r (5 , fNorm , ’ low ’ ) ;

x f = f i l t f i l t ( b1 , a1 , b) ;

%%remove every non−p o s i t i v e value from the data

f o r i= 1 : l ength ( x f )

i f x f ( i )<0

xf ( i ) = 0 ;

end

end

%% Remove only the u s e f u l data and p lo t

xfNew1 = xf (43221 :99472) ; % 180

xfNew2 = xf (467191 :523442) ; %180

xfNew3 = xf (918443 :974693) ; %180

xfNew4 = xf (1374069 :1430319) ; %180

% xfNew5 = xf (3369698 :3425949) ; %180

t =0 :0 . 0032 : 180 ; % Divide the time i n t e r v a l on the time s c a l e by

the d i f f e r e n c e in the index

f o r i= 1 : l ength ( xfNew1 )

i f xfNew1 ( i )<0

xfNew1 ( i ) = 0 ;

end

end

144


p l o t ( t , xfNew1 ( 1 : l ength ( t ) ) ) ; % Equate the dimensions o f the

two parameters .

%Use l ength ( xfNew1 ) , l ength ( t )

to conf i rm

%in the command window

% i f t i s l e s s you p lo t

o therwi se you subt rac t

% from t as p l o t ( t2 ( 1 : l ength ( t2 )

−1) , xfNew2 )

hold on

t2 =180:0 .0032:180+180; % Divide the time i n t e r v a l by the

d i f f e r e n c e in the index


i f xfNew2 ( i )<0

xfNew2 ( i ) = 0 ;

end

end

p lo t ( t2 , xfNew2 ( 1 : l ength ( t2 ) ) ) ; % Equate the dimensions o f the

two parameters .

%Use l ength ( xfNew1 ) , l ength ( t )

to conf i rm

%in the command window

% i f t i s l e s s you p lo t

o therwi se you subt rac t

% from t as p l o t ( t2 ( 1 : l ength ( t2 )

−1) , xfNew2 )

t3= 180+180:0.0032:180+180+180; % Divide the time i n t e r v a l by

the d i f f e r e n c e in the index


145


i f xfNew3 ( i )<0

xfNew3 ( i ) = 0 ;

end

end

p lo t ( t3 , xfNew3 ( 1 : l ength ( t3 ) ) ) ;

t4= 180+180+180:0.0032:180+180+180+180; % Divide the time

i n t e r v a l by the d i f f e r e n c e in the index


i f xfNew4 ( i )<0

xfNew4 ( i ) = 0 ;

end

end

p lo t ( t4 , xfNew4 ( 1 : l ength ( t4 ) ) ) ;

hold o f f

g r i d on ;

t i t l e ( ’ Thrust performance ’ )

x l ab e l ( ’ Time ( s ) ’ ) ; y l ab e l ( ’ Thrust (N) ’ ) ;

%% %%Find the t o t a l impulse o f the response

Tota l Impulse = trapz ( t , xfNew1 ( 1 : l ength ( t ) ) ) + trapz ( t2 , xfNew2

( 1 : l ength ( t2 ) ) ) + trapz ( t3 , xfNew3 ( 1 : l ength ( t3 ) ) ) + trapz (

t4 , xfNew4 ( 1 : l ength ( t4 ) ) )

I 1= trapz ( t , xfNew1 ( 1 : l ength ( t ) ) )

I 2= trapz ( t2 , xfNew2 ( 1 : l ength ( t2 ) ) )

I 3=trapz ( t3 , xfNew3 ( 1 : l ength ( t3 ) ) )

I 4=trapz ( t4 , xfNew4 ( 1 : l ength ( t4 ) ) )

146

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